Liquid Rocket Propellants

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Encyclopedia of Physical Science and Technology EN008B-385 June 29, 2001 15:26

Liquid Rocket PropellantsForrest S. ForbesEngineering Consultant

Peter A. Van SplinterGeneral Physics Corporation

I. Desirable Liquid Propellant CharacteristicsII. Liquid Rocket Propellant Performance

III. Liquid Rocket Propellant SelectionIV. Handling and Storage of Liquid Rocket

Propellants

GLOSSARY

Anhydrous Containing no water. In a practical sense,refers to propellants that have a moisture content belowa specified limit.

Bipropellant Two separate propellants, a fuel and an ox-idizer. They are fed individually to a thrust chamberwhere mixing and combustion occurs.

Cryogenic Liquefied gases that must be refrigerated tobe kept in an liquid state at atmospheric pressure. Noamount of pressure will liquefy gases that are abovetheir critical temperature. An arbitrary temperature of−240◦F (122 K) is sometimes used to define the upperlimit of the critical temperature of cryogenic gases.

Hypergolic Self-igniting. In the context of rocket propul-sion, pertains to fuels and oxidizers that self-ignitewithin about 75 milliseconds after contact with eachother.

MAF Mixed amine fuel. A fuel consisting of mixedamine, usually hydrazine in order to reduce the freezingpoint.

MHF Mixed hydrazine fuel. A special case of MAFwhere only hydrazines are contained.

MON Mixed oxides of nitrogen. An oxidizer consistingof nitrogen tetroxide (N2O4) and nitric oxide (NO). An

associated number represents the NO percentage byweight (e.g., MON-15 contains 15% NO).

Monopropellant Single substance that may be a singlecompound, or a mixture of several materials that arestable under normal conditions, that will combust or de-compose yielding hot gases when ignited or catalyzed.

Neat Liquid propellant composed of only one chemical.Neat hydrazine is only pure N2H4 within the specifica-tions for purity.

Outage Volume of propellant that still remains in the tankafter all usable amounts have been expelled. Volume isretained to assure that the engine is not fed slugs ofpressurizing gas instead of pure propellant.

Passivation Process to render metal and plastic surfacesnonreactive with the propellant.

Prepackaged Rocket system that is loaded and sealedwith storable propellants at the manufacturing plant,thus requiring no servicing of propellants in thefield.

Propellant Fuels, oxidizers, and monopropellants usedto provide energy and the working fluid for the thrustof a rocket engine.

Pyrophoric Spontaneously ignitable in air.Space storable Ability to be used in a space environ-

ment for extended periods without boil-off, freezing,

741

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742 Liquid Rocket Propellants

or decomposition. Usually refers to propellants that aremildly cryogenic or have high vapor pressure on earth,but because of low ambient temperatures in space theycan be kept liquid under a modest vapor pressure. Ex-amples are ammonia, oxygen difluoride, diborane, anddinitrogen tetrafluoride.

Storable Propellants that exhibit no corrosion, decom-position, or deterioration over a specified time period,usually 5–10 years, with 20 years becoming a require-ment for some satellites. A low freezing point is oftenrequired, as low as −65◦F (220 K) for tactical mis-siles, but as high as 30◦F (272 K) for the Titan II.The upper temperature range can also vary, from 90◦F(305 K) for an environmentally controlled system to160◦F (344 K) for tactical missiles. The term storableis often applied to a propellant combination of fuel andoxidizer. While kerosine is storable, it is usually usedwith liquefied oxygen, thus the kerosine/oxygen sys-tem is not storable.

Threshold limit values (TLV) As established by theAmerican Conference of Governmental Industrial Hy-gienists, the time-weighed average concentration of atoxicant to which an average healthy person may berepeatedly exposed for a normal 8-hr day and a 40-hrworkweek without suffering adverse effects. TLV-C isthe concentration that should never be exceeded.

Ullage Volume of a full tank of propellant that is not actu-ally occupied by liquid propellant. The empty volumeallows thermal expansion of the liquid and provides aninitial volume for pressurization gas.

LIQUID ROCKET PROPELLANTS provide the energyor motive power for chemical rocket propelled systemsby releasing chemical energy through a combustion pro-cess. The propellants also furnish the working fluid thatis expanded by the release of the thermal energy. Besidesproducing thrust for prime propulsion or attitude control,propellants are often burned in gas generators to drive tur-bopumps, hydraulic pumps, and alternators. Propellantsare also used to generate gases in underwater flotationdevices, and as the energy source for gas dynamic lasers.

The term liquid rocket propellants encompasses thefuels and oxidizers used in bipropellant systems, andsingle fluid monopropellants. Tripropellant combinationshave also been studied where one component may be asolid. Hybrids are bipropellant systems that have one liq-uid component, usually the oxidizer with the fuel being asolid grain. So called reverse hybrids have liquid fuel and asolid oxidizer grain. Heterogeneous propellants consist offinely powdered energetic materials suspended in a liquidmedium thickened to inhibit settling of the solid particles.

Propellant selection is based on many factors. The idealpropellant does not exist: the large number of propellant

candidates results from the need to make difficult trade-offs among performance, storability, physical character-istics, handling, safety, cost, and availability. All theseparameters are considerations of specific applications andthe degree of importance can vary greatly. A large varietyof liquid fuels, oxidizers, and monopropellants have beenevaluated, some of which are in operational use and othersthat have fallen to the wayside for various reasons.

Synthesis programs have been undertaken to producenew compounds that, theoretically at least, appeared tohave desirable characteristics. Propellant mixtures havebeen formulated in efforts to tailor properties to meetspecific requirements. Propellants are reactive and ener-getic materials by nature. Through an understanding ofthe chemical and physical properties, proper selection ofmaterials and components, and use of appropriate proce-dures and equipment, liquid rocket propellants can be usedsafely and effectively.

I. DESIRABLE LIQUID PROPELLANTCHARACTERISTICS

The primary consideration for propellant selection is highperformance coupled with low system weight. In addition,the physical and chemical characteristics of the propellantsmust meet the requirements of the expected operationalenvironment of the propulsion system. The need for in-stant readiness by the military imposes demands for long-term storability and ease of handling. With the greater useof rocket propulsion for civilian purposes, these consid-erations are less important. Following are the principlefactors.

CHEMICAL ENERGY. Since propulsive energy is pro-duced in the form of hot gases, optimized performance re-sults from maximizing the flame temperature and volumeof gases produced per unit weight of the reactants. A lowmolecular weight of the exit gases is desired.

LIQUID RANGE. The operational environment of themissile or spacecraft dictates the freezing and boilingpoint requirements. Military tactical weapons often re-quire a temperature range of −65–160◦F (219.3–344.4 K)for worldwide deployment. Silo-based missiles and satel-lites may experience a much narrower range. Low-boilingpropellants require heavyweight tanks or some means tokeep the propellants cold. The high vapor pressure compli-cates the design of pump and regeneratively cooled thrustchambers. Special handling procedures and equipment areneeded for servicing cryogenic systems. The freezing andboiling points of various propellants are listed in Tables Iand II. Vapor pressures of selected oxidizers and fuels areplotted in Figs. 1 and 2.

STABILITY. Propellants should be storable for the re-quired time period without decomposition, gas formation,

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Liquid Rocket Propellants 743

TABLE I Physical Properties of Liquid Oxidizers

Freezing point Normal boiling pointa

MolecularOxidizer (formula) weight ◦F K ◦F K Specific gravity

Bromine pentafluoride (BrF5) 174.9 −78.3 211.8 104.9 313.6 2.47 @ 298 K

Chlorine pentafluoride (ClF5) 130.4 −153.4 170.2 7.3 259.4 1.78 @ 298 K

Chlorine trifluoride (ClF3) 92.4 −107.1 195.9 53.2 284.9 1.81 @ 298 K

Florox (CIF3O) 108.4 −86.8 207.1 84.7 302.6 1.85 @ 298 K

Fluorine (F2) 38.0 −362.3 54.1 −306.6 85.0 1.50 @ NBP

Fluorine/Oxygen (Flox) 36.2 — −304 86.5 1.24 @ NBP(70% F2 /30% O2)

Hydrogen peroxide 32.4 11.3 261.6 286.1 414.3 1.39 @ 293 K(90 % H2O2 /(10% H2O)

Hydrogen peroxide 34.0 31.2 272.7 302.4 423.4 1.45 @ 293 K(98% H2O2)

MON-10 (10% NO/90% N2O4) 85.8 −9 250.4 50.8 283.6 1.47 @ 273 K

Nitric acid-type IIIAc 59.4 −56 224.3 140 333b 1.55 @ 298 K

Nitric acid-type IVc 55.0 −35 235.9 76.5 297.9b 1.62 @ 298 K

Nitrogen tetroxide (N2O4) 92.0 11.8 261.9 70.1 294.3 1.43 @ 293 K

Nitrogen trifluoride (NF3) 71.0 −341 65.9 −201 143.7 1.54 @ NBP

Oxygen (O2) 32.0 −361.1 54.7 −297.3 90.2 1.15 @ NBP

Oxygen difluoride (OF2) 54.0 −371 49.3 −228.6 128.4 1.52 @ NBP

Ozone (O3) 48.0 −316 79.8 −170 160.9 1.61 @ 78 K

Perchloryl fluoride (ClO3F) 102.4 −231 127 −52.3 226.3 1.39 @ 298 K

Tetrafluorohydrazine (N2F4) 104.0 −261.4 110.2 −99.4 200.2 1.56 @173 K

a NBP is normal boiling point at 1 atm.b Bubble point, where liquid appears to boil.c See Table X for nitric acid compositions.

or attack on the container material. Decomposition orresidue formation at high temperatures can adversely af-fect regenerative cooling. Propellants should also be sta-ble to mechanical and hydrodynamic shock and adiabaticcompression.

REACTIVITY. Mild corrosion of tanks and compo-nents and deterioration of seals and gaskets can resultin propellant contamination and clogging of filters andsmall passages. Reactivity with moisture and air invokesthe need for unique transfer techniques. Spills and leaks ofreactive propellants are safety hazards; benign propellantsare easily and safely handled with a minimum of specialequipment. A reactive fuel and oxidizer combination isneeded for smooth and efficient combustion.

DENSITY. High propellant density allows the use ofsmaller tanks, thus minimizing the overall vehicle inertweight (improves the mass fraction). A propellant withlow coefficient of expansion can be used over a larger por-tion of its liquid range without undue performance vari-ation. The specific gravities of common propellants arepresented in Tables I and II and as a function of tempera-ture in Fig. 3.

VISCOSITY. Low viscosity fluids will have less pres-sure drop through the feed system and thrust chamber

injector, permitting a smaller pump or a lightweight pres-surization system to be used.

IGNITION. Hypergolic ignition offers high reliability, arestart capability for multiburn missions and eliminates theweight of a separate ignition system. In space, secondaryignition sources are unreliable. Catalytic decompositionof monopropellants is a valuable attribute.

LOGISTICS. While low-cost propellants are desirable,they are usually a small part of the overall system cost.The costs of special handling and storage, as in the case ofcryogens, must be considered part of the total propellantcost. Propellant costs are a function of the production rate.Availability of propellant for a new or expanded programmay involve construction of a new plant or expandingan existing one. The required transportation equipment isanother factor.

II. LIQUID ROCKET PROPELLANTPERFORMANCE

A. General Principles

It is helpful to explore the governing equations for rocketengines in order to understand the motivation of rocket

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TABLE II Physical Properties of Liquid Fuels

Freezing point Normal boiling pointMolecular Specific

Fuel (formula) weight ◦F K ◦F K gravity

Aerozine-50 (50% Hydrazine/50% UDMHa) 41.8 22 267.6 158 343.15 0.899 @ 298 K

Ammonia (NH3) 17.0 −107.9 195.4 −28.05 239.8 0.682 @ NBP

Aniline (C6H5–NH2) 93.12 207.4 266.8 363.9 457.5 0.999 @ 293 K

Diborane (B2H6) 27.69 −264.8 108.3 −134.5 180.6 0.438 @ NBP

Diethylcyclohexane [(C2H5)2C6H10] 140.3 −110.2 194.2 345 447.2 0.804 @ 293 K

Diethylenetriamine (DETA) [(NH2C2H4)2NH] 103.2 −38.2 234 405 480 0.953 @ 298 K

u-Dimethyl hydrazine (UDMH) [CH3)2N2H2] 60.1 −70.9 216.0 144.2 335.5 0.791 @ 298 K

Ethane (C2H6) 30.07 −297.9 89.95 −127.6 184.48 0.548 @ NBP

Ethanol (C2H5OH) 46.1 −174 158.7 172.9 351.4 0.789 @ 293 K

Ethene (C2H4) 28.054 −273.1 103.6 154.66 169.4 0.5674 @ NBP

Ethylene dihydrazine (C2H10N4) 90.1 55.4 285.95 >491.e >528.e 1.096 @ 298 K

Ethylene oxide (C2H4O) 44.01 −168 162 50.9 283.6 0.887 @ NBP

Furfuryl alcohol (C5H5OOH) 98.1 −26 240 340 444 1.13 @ 293 K

Hybaline A-5 109.2 −58 223.2 505 536 0.736 @ 293 K

Hydrazine (N2H4) 32.04 34.75 274.7 236.3 386.6 1.008 @ 293 K

Hydrogen (H2) 2.016 −434.8 13.8 −423.3 20.21 0.0709 @ NBP

JP-X (40% UDMH/60% JP-4)a 89.3 <−71 216 211 373 0.80 @ 293 K

Kerosine (RP-1) (H/C = 2.0) 172 <−50 228 350–525 450–547 0.80–0.81

450–547 @ 293 K

Methane (CH4) 16.04 −296.5 90.6 −258.7 111.6 0.451 @ NBP

Methanol (CH3OH) 32.04 −144 175.4 147 337 0.791 @ 293 K

MAF-1b 100.0 −148 173 170 350 0.87 @ 298 K

MAF-3b 90.2 <−65 219 — 0.917 @ 298 K

MHF-3c (86% MMH/14% N2H4)a 43.41 <−65 219 193.4 362.8 0.889 @ 298 K

MHF-5c 45.35 −43.6 231.2 205.8 369.7 1.011 @ 298 K

Monomethyl hydrazine (CH3H2H3) 46.07 −62.3 220.8 189.8 360.8 0.879 @ 293 K

Nitromethane (CH3NO2) 61.04 −20.2 244.1 214 374.3 1.135 @ 298 K

Pentaborane (B5H9) 63.17 −52.2 226.4 140.1 333.2 0.623 @ 298 K

n-Propyl nitrate (C3H7NO3) 105.1 <−150 172 230.9 383.6 1.058 @ 298 K

Otto Fuel IId 186.52 −18.4 245.2 decomposes 1.232 @ 298 K

Propane (C3H8) 44.11 −305.84 85.5 −43.7 231.1 0.5853 @ NBP

Quadricyclane (C7H8) 92.14 226.4 381 0.982

Shelldyne-H 186.7 −101 200 460 511 1.11 @ 293 K

U-DETA (MAF-4, Hydyne) 72.15 <−120 188.7 161 345 0.858 @ 289 K(60% UDMH/40% DETA)a

a Weight percent.b See Table XII for composition.c See Table XIII for composition.d See Table XV for composition.e Estimated values courtesy of GenCorp Aerojet Tech Systems.

engineers to create and use cryogenic, toxic, and explo-sive compounds in the pursuit of rocket-powered vehi-cles with minimum weight or minimum volume (or both).Liquid rockets are reaction engines derived from New-ton’s Laws. Saenger described the momentum princi-ple with these words, “. . . all ship propellers, airplanepropellers, water wheels, and oars generate their for-ward push at the expense of the momentum of wa-

ter or air masses which are accelerated toward therear.”

Therefore, accelerating a working fluid out the rear ofa vehicle produces thrust. The working fluid can be coldgas, monopropellant decomposition products, bipropel-lant combustion products, or molecules that have beenheated by an electrical, solar or nuclear energy source.The thrust of a liquid propellant rocket derives primarily

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FIGURE 1 Vapor pressure of liquid oxidizers.

from a high velocity of exhaust products exiting the engineat the exit plane.

A simple party balloon inflated with air illustrates allthese principles. The energy source is the elastic energyof the stretched rubber. The working fluid is the air en-closed within the inflated balloon. The exhaust nozzle isthe hole where the air can escape. At the exit plane of theballoon, the air is flowing at a higher velocity than thebulk of air still remaining in the balloon. The applicableNewton’s Law is that the time rate of change in momen-tum will be proportional to the net external force. Thisis the first equation, the classic equation for thrust of arocket.

F = mV e + (Pe − Pa)Ae (1)

Where:

m is mass flow rate of working fluid (lb/sec, Kg/s)Ve is velocity of the working fluid where it exits thefree body (ft/sec, m/s)Pe is the pressure of the working fluid at the place theworking fluid exits the free body (lb/ft2, Pascals)

Pa is the pressure in the surroundings of the free body(lb/ft2, Pascals)Ae is the area of the exit of the nozzle (ft2, m2)

From 90 to 100% of the thrust is contributed by the firstterm, and the term may be called velocity thrust or mo-mentum thrust. In this term it can be seen that the higherthe flow rate, the higher the thrust, and the higher the exitvelocity, the higher the thrust. For a fixed flow rate ofworking fluid, this equation shows why rocket engineersattempt to maximize thrust per unit flow rate by maximiz-ing exit velocity.

The second term of the thrust equation commonly con-tributes 0–10% to the thrust and is called “pressure thrust.”Pressure thrust is greatly dependent upon the pressureenvironment surrounding the free body. For instance, atliftoff, the pressure thrust of the Space Shuttle Main En-gine (SSME) is actually negative since Pa exceeds Pe atsea level. In space vacuum, where Pa is practically zero,the pressure thrust of the SSME contributes on the orderof 10% to the thrust. For a vehicle operating at constant al-titude, arranging so that Pe = Pa results in what is defined

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FIGURE 2 Vapor pressure of liquid fuels.

as optimum expansion ratio where pressure thrust is de-liberately set to be zero. Allowing Pe to be greater than Pa

may be thought of as wasted energy that could have beenconverted to more thrust at the same flow rate of propellant

Thrust per unit flow rate is defined as specific impulse(Isp)

Isp = F/m (2)

Specific impulse is the conventional parameter for theoret-ical comparison of various propellants and for assessmentof the efficiency of actual delivered rocket engine perfor-mance compared the the theoretical.

In the English units, m is the weight flow rate in poundsper second. Since the thrust is in pounds, this leaves theunit of specific impulse to be seconds. For a more physicaldefinition, specific impulse must be thought of as poundsof force per pound per second of propellant flow. In anyunits, the higher the value of specific impulse, the higherthe specific performance of the rocket. In the MKS systemof units, if thrust is in newtons and mass flow rate is inkilograms per second, the units of specific impulse willreduce to meters per second.

A propellant combination with higher specific impulsewill result in either a lighter vehicle for a fixed payloador more payload for a fixed weight of vehicle. For mis-siles the trade might be range and payload. The lighter thevehicle, the easier and cheaper it might be to construct,transport, and handle. Furthermore, lighter may mean lessraw material of construction that might translate to lowercost. Vehicle designers tend to choose the highest specificimpulse propellant combination, if all other characteris-tics are equal or at least acceptable. Of course, all othercharacteristics are never equal, so for different kinds of ve-hicles, different propellant combinations provide the bestcompromise.

Returning to Eq. 1, dividing by m and arranging foroptimum expansion ratio (Pe = Pa) the following results:

F/m = V = c (3)

This velocity that remains on the right side of the equationcan be interpreted as the actual physical velocity of theworking fluid relative to the rocket vehicle. This velocityis defined as effective exhaust velocity and is commonlyassigned the letter c. Note that effective exhaust velocity

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FIGURE 3 Specific gravities of liquid rocket propellants.

in the metric system is also the specific impulse value inthe metric system for optimum expansion ratio. When arocket is reported to have a specific impulse of 1000 m/sec,it means that the velocity of gases exiting the nozzle is1000 m/sec relative to the nozzle.

In the air-filled party balloon, the air is a compressiblefluid being accelerated out of an exit hole. For a rocketengine, the exit hole is the exit plane of a converging-diverging supersonic nozzle. From the principles of ther-modynamics and fluid mechanics, an equation can bederived using conservation of energy, perfect gas, and aconcept called a reversible adiabatic flow. Without the de-tails of a derivation, the velocity out of the exit of a nozzleis written as:

Ve =√√√√ k

k − 12g

R

MWTc

[1 −

(Pe

Pc

) k−1k

](4)

Where:

Ve is exit velocity of the working fluidk is the ratio of specific heats of the working fluidg is the gravitational constantR is the universal gas constant

MW is the molecular weight of the working fluidTc is the absolute temperature of the working fluid atthe entrance of the nozzlePe is the pressure of the working fluid at the exit ofthe nozzlePc is the pressure of the working fluid at the entranceof the nozzle

Approximately, Pc and Tc are the pressure and temper-ature of the working fluid in the combustion chamber.

From Eq. 4 several of the driving factors for propel-lant selection can be recognized. To maximize Ve whichmaximizes thrust, the desirable factors are

1. The lowest value of k2. The lowest value of molecular weight3. The highest value of Tc

4. The highest value of Pc

The value of k varies from 1.28 for helium to 1.66 forcarbon dioxide. Air, hydrogen, oxygen, and nitrogen are1.4. This value is not really selectable, however, the lowestmolecular weight gases also tend to have the lowest k.

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The lowest molecular weight is 2.02 for hydrogen. Hereis the driver for the selection of liquid hydrogen, hydrides,and all other fuels packed with as many hydrogen atomsas science can achieve. Carrying along nitrogen, oxygen,and carbon compounds in the exhaust increases the mix-ture molecular weight and penalizes Ve. If the source ofheat energy is from a solar concentrator, a nuclear reactor,or electrical energy, the working fluid can be pure hy-drogen. However, when the source of energy is chemicaldecomposition or combustion, this increased molecularweight of the working fluid is unavoidable.

For a chemical rocket, energy released in the combus-tion chamber heats the working fluid to very high tempera-tures. Maximum combustion chamber temperature yieldsmaximun exit velocity. Here is the driver for selection ofoxidizers of legendary reactivity in an attempt to run Tc ashigh as possible while minimizing melting all the parts ofthe rocket. Some of the highest heat releases stem from theoxidation of metals. This is the driver for including metal-lic aluminum or even beryllium as part of the combustionprocess.

High chamber pressure, Pc, is a factor that could allowenormous expansion ratios that still result in Pe greaterthan Pa . When Pe is greater then Pa , it might be considereda loss of opportunity to convert that extra pressure to extraexhaust velocity. Throwing away the extra exit pressuremeans that specific impulse could have been higher.

To summarize, the vehicle designer wants the lowestmolecular weight exhaust, the highest combustion tem-perature, and the highest chamber pressure that the themission requires or the sponsor is willing to pay for, andit is all caused by the relationship of Eq. 4.

Returning to Eq. 1, if mass flow is replaced by the con-tinuity equation, velocity is replaced by Eq. 4, the perfectgas equation of state is applied, and variables are mathe-matically manipulated, the thrust equation becomes:

F = Pc At C f (5)

where Pc is combustion chamber pressure, At is the areaof throat of the converging-diverging supersonic nozzle,and Cf is a complicated term involving k, Pc, and Pe andis called the thrust coefficient.

Equation 5 is an interesting way of arriving at rocketthrust because it separates the pressure thrust (pressuretimes area) of the rocket chamber from the contribution ofthe converging-diverging supersonic nozzle. All the messymethematics is in Cf . Values of Cf vary from approxi-mately 1.0 to a maximum theoretical value of 1.964 forinfinite ratio of Pa /Pe. So the thrust of an actual rocket issomewhere between 100 and 200% of the pressure thrust,depending upon the nozzle expansion ratio.

If F is replaced by Eq. 5 in the definition of specificimpulse (Eq. 2), then:

Isp = Pc At C f

m(6)

A new term can be defined by collecting all the vari-ables except thrust coefficient. That new variable is de-fined as the characteristic exhaust velocity, called c∗ andpronounced “see star.” Physically, it may be thought ofas the velocity of the working fluid at the throat of theconverging-diverging nozzle.

c∗ = Pc At

m(7)

Replacing Eq. 7 in Eq. 6 shows an interesting relationshipfor Isp:

Isp = c∗C f (8)

Again, the performance of the nozzle has been separatedfrom the performance of the chamber. In this way, c∗ isrepresentative of the combustion properties of the propel-lants used in the rocketwhile C f is representative of thenozzle.

Tests can be run at atmospheric pressure without thetrouble of measuring thrust in order to optimize mixtureratio of a bipropellant or to evaluate the survivability ofthe hardware to the effects of high Pc and Tc. It is mucheasier and faster to accurately measure Pc of prototype,heavyweight rocket hardware than it is to measure thrust.As a result, the first measure of how well an injector per-forms is nearly always c∗ efficiency which is measured c∗

compared to theoretical c∗.The above relationships will predict or measure the per-

formance of rocket engines within 10%. Of course, the realworld will impose the effect of real gases, three dimen-sions, and various other losses on the process of devel-opment, but the basic selection and evaluation parametersfor propellant definition have been defined.

B. Theoretical Calculations

Specific impulse is also defined from the total enthalpychange of the propulsion system between the combustionchamber and the nozzle exit:

Isp =√

2J (Hc − He)/g (9)

where Hc is enthalpy of combustion products before ex-pansion, Btu/lbm (J/kg), He enthalpy of combustion prod-ucts after expansion, Btu/lbm (J/kg), and J mechanicalequivalent of heat 778 ft lbm/Btu (0.101988 kg m/J). Twosets of conditions are considered when making perfor-mance calculations, those of the combustion chamber andthose of the expansion nozzle.

1. Combustion Chamber

The conditions of the propellant entering into the combus-tion chamber (composition, temperature, physical state,

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and mixture ratio) and the chamber pressure are used tocalculate the chamber temperature and hot gas reaction,product composition, and properties. This information de-fines the nozzle inlet conditions that allow calculation ofIsp and C ∗ and provides information needed for enginedesign, estimates of heat transfer, and nozzle flow and ex-pansion parameters. The combustion process is assumedto be adiabatic and proceeds to completion.

Chamber (flame) temperature, Tc, and gas compositionare calculated using mass, pressure and energy balanceequations, and chemical equilibria data.

The energy equation does not provide a direct solutionof the chamber temperature. When the enthalpy of thereaction products above a standard reference tempera-ture equals the heat of reaction of the entering propellant,the adiabatic combustion temperature is defined. For ma-chine computation, enthalpy values of the products arerepresented as a power series in terms of the temperature.The general energy equation is∑i,propellant

ni{[(

HT◦c − H ◦0

) − (H ◦T0

− H ◦0)] + (

�H ◦f)

T 0

}i

=∑

j,reaction products

nj{[(

H ◦Ti− H ◦0

) − (H ◦T0

− H ◦0)]

+ (�Hf δ)T0

}j (10)

where Tc is adiabatic flame combustion temperature, Ti

propellant temperature, T0 reference temperature (usually

TABLE III Thermodynamic Properties of Liquid Oxidizers

Critical Heat ofHeat of temperature Critical pressure vaporization

formationOxidizer (kJ/mol)a ◦F K psia MN/m2 Btu/lb kJ/mol

Bromine pentafluoride (BrF5) −548.94 388 471 669 4.612 74.1 30.12

Chlorine pentafluoride (ClF5) −253.13 289.4 416.2 771 5.316 73.3 22.22

Chlorine trifluoride (ClF3) −185.77 355 453 961 6.626 128.0 27.53

Florox (ClF3O) −165.69 360 455 — 127.9 32.22

Fluorine (F2) −12.96 −199.6 144.5 809.7 5.583 71.5 6.32

Fluorine/oxygen (70/30) (FLOX) −36.04 −199.6 144.5 809.7 5.583 77 6.48

Hydrogen peroxide (H2O2) −187.78 855 730 3130 21.58 596 47.07

Nitric acid-Type IIIAb −140.16 520 544 1286 8.867 247 34.10

Nitric acid-Type IVb −181.59 512 540 1428 9.846 270 34.52

Nitrogen tetroxide (N2O4) −19.58 316.8 431.4 1441 9.938 178.2 38.12

Nitrogen trifluoride (NF3) −131.50g −88 206.5 729 5.026 73.0 11.59

Oxygen (O2) −12.12 −181.1 154.8 730.6 5.375 91.6 6.82

Oxygen difluoride (OF2) +24.52 −72 215.4 719 4.958 88.7 11.13

Ozone (O3) +118.41 9.6 260.7 803 5.537 127.9 14.27

Perchloryl fluoride (ClO3F) −21.42 203 368 779 5.372 83.9 19.96

Tetrafluorohydrazine (N2F4) −21.8 96.8 309 1130 7.8 64.2 15.23

a Liquid state at 298.16 K. To obtain Kcal/mol divide kJ/mol by 4.184.b See Table X for nitric acid compositions.

298 K), H ◦ - H ◦0 molar enthalpy (from tables), �H ◦f heatof formation, ni moles of reactants (propellants), and nj

moles of reaction products. To obtain values for the un-knowns, Tc and nj, the equations are solved simultane-ously with a digital computer using thermochemical andequilibria data from references such as the JANAF Ther-mochemical Tables and the National Bureau of StandardsCirculars. These tabulations cover heat capacity, enthalpy,entropy, free energy function, heat and free energy offormation, and equilibrium constants of formation overthe range of 0–6000 K, as appropriate. Heats of forma-tion of selected propellants are presented in Tables IIIand IV.

2. Expansion Nozzle

In the nozzle, the combustion gases are expanded isen-tropically, converting sensible enthalpy into kinetic en-ergy accompanied by a drop in temperature and pressure.When the gas composition does not change (no chemicaldissociation, recombination, or condensation) from thechamber to the nozzle exit plane, the process is consid-ered to be in frozen equilibrium. If chemical and phaseequilibrium among all combustion species is maintainedunder the varying pressure and temperature conditions ofthe nozzle expansion process, the product compositionwill change. This is known as shifting equilibrium, andperformance so calculated will bethe higher value because

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TABLE IV Thermodynamic Properties of Liquid Fuels

Critical Critical Heat ofHeat of temperature pressure vaporization

formationFuel (formula) (kJ/mol)a ◦F K psia MN/m2 Btu/lb kJ/mol

Aerozine-50 (50% hydrazine/50% UDMH)b +51.51 633 607 1731 11.935 346.5 33.67

Ammonia (NH3) −71.71 270.1 405.4 1636 11.280 596.2 23.56

Aniline (C6H5–NH2) +30.71 798 698.6 771 5.316 203 43.93

Diborane (B2H6) +20.79 62.1 289.8 581 4.006 224.3 14.43

Diethylcyclohexane [(C2H5)2C6H10] −726.18 690 638.6 367.5 2.534 119 38.79

Diethylenetriamine (DETA) [(NH2C2H4)2NH] −77.40 433.9 496 538 3.710 210.5 50.50

u-Dimethyl hydrazine (UDMH) [(CH3)2N2H2] +51.63 482 523 867 5.978 250.6 35.02

Ethane (C2H6) −25.26 90.1 305.45 708.5 4.885 210.0 14.67

Ethanol (C2H5OH) −277.65 469.4 516 925 6.378 362 38.79

Ethene (C2H4) 8.1 49.82 282.7 742.35 5.118 207.5 13.55

Ethylene dihydrazine (C2H10N4) 137.7 901. f 746. f 838. f 5.8 f 367.7 f 77.0 f

Ethylene oxide (C2H4O) +25.15 384.4 468.9 1043 7.191 249 25.48

Furfuryl alcohol (C5H5OOH) −276.35 570 f 572 f 515 f 3.5 f 235.2 53.62

Hybaline A-5 +46.02 662–932 623–773 — 82.26 20.88

Hydrazine (N2H4) +50.42 716 653 2131 14.693 583 43.43

Hydrogen (H2) −9.01 −399.9 33.21 188 1.296 195.3 0.915

JP-X (40% UDMH/60% JP-4)b +4.90 545 558 680 4.689 169 35.06

Kerosine (RP-1) (H/C = 2.0) −26.03 758 676.5 315 2.172 125 49.79

Methane (CH4) −89.50 −115.8 191.0 673 4.64 238.9 8.91

Methanol (CH3OH) −239.03 464 513 1142 7.87 526.8 39.32

MAF-1c −77.82 — — 246 56.90

MAF-3c −38.66 — — —

MHF-3d (86% MMH/14% N2H4)b +54.00 617 598 1373 9.467 370.2 37.38

MHF-5d +28.75 685 636 1470 10.14 324.2 34.18

Monomethyl hydrazine (MMH) (CH3N2H3) +54.84 594 585 1195 8.24 377 40.38

Nitromethane (CH3NO2) −139.03 598.6 587.9 916 6.314 269.8 38.28

Pentaborane (B5H9) +32.38 435 497 557 3.84 219 32.17

n-Propyl nitrate (C3H7NO3) −214.47 584.6 580 588 4.05 142 34.68

Otto Fuel IIe −410.16 (decomposes) —

Propane (C3H8) −30.372 206.2 369.93 617.4 4.257 183.2 18.78

Shelldyne-H (RJ-5) +136.82 985 803 510 3.52 191 82.88

U-DETA (MAF-4, Hydyne) (60% UDMH/ +16.44 558 565.4 805 5.55 191 32.0140% DETA)b

a Liquid state at 298.16 K. To obtain Kcal/mol divide kJ/mol by 4.184.b Weight percent.c See Table XII for composition.d See Table XIII for composition.e See Table XV for composition.f Estimated values courtesy of GenCorp Aerojet TechSystems.

the dissociated species that recombine in the nozzle addto the release of chemical energy converted into kineticenergy. The frozen and shifting equilibrium representthe high and low limits of performance due to nozzlechemistry attainable from the system. In practice, actualrocket performance usually falls in between the two. Otherlosses, such as incomplete combustion, heat loss, noz-

zle friction, and divergence angle may play a significantrole.

To perform frozen equilibrium calculations, the exittemperature, Te, is determined by considering the conser-vation of entropy (an isentropic process). The gas compo-sition throughout the nozzle is the same as in the chamber,thus the entropy in the chamber and at the nozzle exit are

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equal. The difference between frozen and shifting flowvalues occurs because the species mole fractions areknown for the former and unknown for the latter, whichmakes the latter an iterative process similar in complexityto the chamber solution. Usually when making a frozenflow calculation an additional simplifying assumption ismade: that the heat capacity is constant with temperature.This leads to a simple expression for the effect of temper-ature on entropy:

S = Cp ln(T2 /T1) (11)

This entropy change is set equal and opposite to the changedue to the pressure drop:

S = −R ln(T2 /T1) (12)

leading to the familiar isentropic relationship:

ln(T2 /T1) = R /Cp ln(P2 /P1) = (k − 1)/k ln(P2 /P1)(13)

where

k = Cp /Cv = Cp /(Cp − R) (14)

Perhaps a yet more familiar form is:

T2 /T1 = (P2 /P1)(k −1)/k (15)

For a selected exit pressure, the exhaust gas temperaturecan now be calculated. The change in enthalpy leads tothe calculation of the specific impulse by Eq. 9.

In shifting equilibrium calculations, the gas composi-tion changes as it transverses from the chamber throughthe nozzle to the exit plane. As the temperature and pres-sure drops, many of the dissociated species recombinewhich change the gas composition and release additionalenergy. The mass balance, pressure balance, and chemi-cal equilibrium must be satisfied as described previouslyfor the chamber conditions, and the exit temperature, Te,calculated. However, instead of the energy balance Eq. 10an expression of entropy conservation is used to solve forthe exit temperature and the exhaust composition:∑

k,exhaust

nk[STe − R ln p

]k =

∑j,chamber

nj[STc − R ln p

]j

(16)

where Te is exhaust temperature at exist plane, Tc cham-ber temperature, STc absolute molar entropy of speciesj at chamber temperature and 1 atm, STe absolute molarentropy of species k at exit temperature and 1 atm, R gasconstant, nk number of moles of species k at nozzle exit, nj

number of moles of species j from chamber, and pj,k partialpressures (atm) of components. From the temperature andproduct composition, one can calculate the enthalpy dropin the nozzle, Isp (shifting) is then calculated from Eq. 9.Refinement of the thermochemical performance calcula-tions can be made to correct for two-phase losses when

condensed species are present, and for friction and heattransfer losses. Also, for an equilibrium flow system inwhich the nozzle species and average molecular weightare changing, the optimum nozzle area ratio for a givenexit pressure can be calculated.

A familiar expression of specific impulse can be derivedfrom Eqs. 9 and 15 and H = CpT :

Isp =√√√√ k

k − 1

2RTc

gm

[1 −

(Pe

Pc

) k −1k

](17)

where k is the ratio of specific heats (Cp /Cv), Pe exitpressure, Pc chamber pressure, Tc chamber temperature,R universal gas constant, and M average molecular weightof combustion products. As noted in Eq. 17 Isp variesdirectly with the square root of the chamber temperatureand inversely with the square root of the average molecularweight of the exhaust products, thus the ratio (Tc/M)1/2

is used as a rough figure of merit to compare propellants.The theoretical performance of selected liquid propel-

lant systems is presented in Tables V–IX.

III. LIQUID ROCKET PROPELLANTSELECTION

The best oxidizer performance is provided by the el-ements O and F. Unfortunately, their molecular forms,O2 and F2, are cryogenic. The elements Cl and Br willgive oxidizers better physical properties, but their highmolecular weight makes them less attractive from a per-formance standpoint. Nitrogen serves as a “back-bone”to carry O and F in a number of oxidizers, but adds lit-tle to the overall performance. Although hydrogen is afuel, it is an excellent working fluid because of its lowmolecular weight; thus, it is a desirable component insome oxidizers. From these elements, the following ox-idizers are obtained: O3, F2, OF2, NF3, O2, N2F4, ClF5,ClO3F, ClF3O, ClF3, N2O4, H2O2, HNO3. This listing isfrom the highest performance to the lowest using hydrogenas the fuel. (With another fuel, some reordering may oc-cur.) The first five oxidizers are the highest energy and arecryogens.

The number of elements that can be considered as fuelsor fuel components is limited. Based on heat release, orheat of combustion, the elements H, Li, Be, B, C, Mg, andAl can be considered. As in the case with oxidizers, ni-trogen will contribute little to the heat of combustion butserves as a desirable building block, especially for hold-ing hydrogen to produce the amine group. Carbon alsois a useful hydrogen carrier, leading to the hydrocarbonfamily. The variety of liquid compounds that can be madefrom these elements is a further restriction. Metals are

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TABLE V Theoretical Performance of Rocket Liquid Bipropellant Combinations (Cryogenic Systems)a

Specific impulse (sec) Combustion temperatureBulk O/F

Oxidizer Fuel density ratiob Frozenc Shiftingc ◦F K

Oxygen Hydrogen 0.28 4.02 387 391 4935 2997

Oxygen Kerosine 1.02 2.06 285 300 6160 3677

Oxygen u-Dimethyl hydrazine 0.98 1.65 295 310 6010 3594

Ozone Hydrogen 0.30 4.0 413 424 5523 3323

Fluorine Hydrazine 1.31 2.30 334 363 7955 4675

Fluorine Hydrogen 0.45 7.60 396 410 6505 3869

Oxygen difluoride Hydrazine 1.27 1.54 317 339 6690 3972

Oxygen difluoride Diborane 1.00 3.73 340 365 7964 4680

Flox-30d Kerosine 1.09 3.05 298 316 6623 3935

Flox-70e Kerosine 1.20 3.80 317 344 7820 4600

a Courtesy of Rocketdyne, Division of Rockwell International and the U.S. Air Force Rocket Propulsion Laboratory.b Oxidizer to fuel mixture ratio by weight.c Pc = 1000 psia expanded to 14.7 psia (6895 kN/m2 → 101.4 kN/m2).d A mixture of 30% fluorine and 70% oxygen, by weight.e A mixture of 70% fluorine and 30% oxygen, by weight.

solids, of course, as are most of their hydrides. Metals donot have a heat release appreciably greater than hydrogenand, with oxygen-based oxidizers, solid oxide productsform that cannot be expanded through the nozzle, causinga loss of specific impulse. From these considerations, weobtain the following fuel categories: hydrogen, ammoniaand derivatives (amines, hydrazines), hydrocarbons, boro-hydrides, and alcohols.

TABLE VI Theoretical Performance of Rocket Liquid Bipropellant Combinations (Storable Systems)a

Specific impulse (sec) Combustion temperature

Oxidizer Fuel Bulk density O/F ratio Frozenb Shiftingb ◦F K

Tetrafluorohydrazine Ammonia 1.27 4.49 307 325 7284 4302

Tetrafluorohydrazine Hydrazine 1.43 3.16 313 334 7547 4448

Chlorine trifluoride Hydrazine 1.50 2.78 279 295 6553 3896

Hydrogen peroxide (95%) Hydrazine 1.26 2.00 269 287 4814 2657

Hydrogen peroxide (95%) Pentaborane 1.04 1.06 299 308 5000 3033

Hydrogen peroxide (95%) Kerosine 1.30 7.35 265 273 4785 2914

Chlorine pentafluoride Hydrazine 1.47 2.66 294 313 7039 4166

Chlorine pentafluoride MHF-3 1.41 2.79 276 303 6470 3850

Chlorine pentafluoride NOTSGEL-A 1.70 2.6 262 285 8058 4732

Nitrogen tetroxide Monomethyl hydrazine 1.20 2.15 278 289 5653 3396

Nitrogen tetroxide Hydrazine 1.22 1.30 284 292 5406 3259

Nitrogen tetroxide Aerozine-50c 1.21 2.00 276 289 5590 3361

Nitrogen tetroxide Aerozine-67e 1.28 1.80 273 288 6094 3368

MON-25 Monomethyl hydrazine 1.18 2.28 279 290 5705 3425

IRFNA, Type IIIA u-dimethylhydrazine 1.26 2.99 267 276 5300 3200

Hydrazine Pentaborane 0.79 1.28 325 328d 4661 2845

a Courtesy of Rocketdyne, Division of Rockwell International and the U.S. Air Force Rocket Propulsion Laboratory.b Pc = 1000 psia expanded to 14.7 psia (6895 kN/m2 → 101.4 kN/m2).c A mixture of 50% hydrazine and 50% UDMH, by weight.d When kinetic limitations are considered, Isp = 315 seconds.e A mixture of 67.2% ethylene dihydrazine and 32.8% hydrazine, by weight.

A. Cryogenic Oxidizers

1. Ozone

Liquefied ozone, O3, has stirred the interest of propellantchemists since the late 1940s. It has a serious disadvantagein that with very little stimulus it reverts back to oxygen ex-plosively. Many impurities are known to sensitize ozone,such as hydrocarbons, nitrogen oxides, silver, and copper.

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TABLE VII Theoretical Performance of Rocket Liquid Bipropellant Combinations (Space Conditions—Vacuum Expansion)a


Oxidizer Fuel Bulk density O/F ratiob Frozenc Shiftingc ◦F K

Oxygen Hydrogen 0.31 4.50 446 456 5012 3040

Oxygen Methane 0.82 3.32 339 365 5482 3301

Oxygen difluoride Diborane 1.00 3.82 380 428 7054 4174

Chlorine pentafluoride Hydrazine 1.47 2.72 321 363 6533 3885

Chlorine pentafluoride MHF-3 1.41 2.8 305 354 6010 3594

Fluorine Hydrazine 1.31 2.34 369 424 7298 4310

Fluorine Hydrogen 0.50 9.00 450 475 6388 3804

Tetrafluorohydrazine Hydrazine 1.43 3.24 348 387 6978 4132

Nitrogen tetroxide Monomethyl hydrazine 1.21 2.28 317 340 5300 3200

Nitrogen tetroxide Aerozine-67 1.28 1.85 307 338 5741 3172

Nitrogen tetroxide Hybaline A-5 1.22 2.33 — 380 6596 3920

a Courtesy of Rocketdyne, Division of Rockwell International, and the U.S. Air Force Rocket Propulsion Laboratory.b Oxidizer to fuel mixture ratio by weight.c Pc = 150 psia (1034 kN/m2) expanded to Ae/At = 40.

The U.S. Navy and Air Force funded considerable researchstudying ozone decomposition, effect of additives (sta-bilizers and impurities), and analytical techniques. Somefluorine compounds tended to improve ozone stability, butnot sufficiently to make ozone a useful rocket oxidizer. Fir-ings of a 25% O3/75% O2 mixture were accomplished atthe 100-lb thrust (445 N) level using alcohol and kerosineas fuels.

2. Fluorine

Liquefied fluorine (F2) has been extensively evaluatedand considerable rocket engine development has beenaccomplished at the 15,000 and 40,000 lb f (66.7 and177.9 kN) level, although it has not yet been selected for

TABLE VIII Theoretical Performance of Rocket Liquid Tripropellant Combinationsa


Oxidizer Fuelb Bulk density O/F ratio Frozen Shifting ◦F K

Oxygen Hydrogen (50%) 0.23 0.88 453c 458c 4662 2845

Beryllium (50%)Oxygen Hydrogen (49%) 0.24 0.92 513 539d 4681 2856

Beryllium (51%)Fluorine Hydrogen (69%) 0.17 0.85 425c 432c 3056 1953

Lithium (31%)Fluorine Hydrogen (68%) 0.17 0.88 468 506d 2905 1869

Lithium (32%)Hydrogen peroxide Hydrazine (71%) 1.24 0.55 329c 334c 5606 3370

Beryllium (29%)

a Courtesy of Rocketdyne, Division of Rockwell International, and the U.S. Air Force Rocket Propulsion Laboratory.b Percents are by weight.c Pc = 1000 psia expanded to 14.7 psia (6895 kN/m2 → 101.4 kN/m2).d Pc = 150 psia expanded to Ae/At = 40.

an operational system. A very reactive material, fluorinewill react with most anything, thus all fluorine systemsmust be scrupulously cleaned and passivated. Mixtures offluorine and oxygen (Flox) have been evaluated as a meansof increasing oxygen performance. The experimental ef-forts have indicated that mixtures containing 30% or lessfluorine behave essentially like oxygen, thus Flox30 couldbe substituted for liquefied oxygen with minimal changesfor about a 5–10% performance gain, depending on thefuel selected.

3. Oxygen Fluorides

Liquefied oxygen difluoride (OF2) has been of interestfor its “space storability” aspect. Its boiling point,

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TABLE IX Theoretical Performance of Rocket Liquid Monopropellantsa

Specific impulse (sec) Combustion temperatureChamber

Monopropellant pressure (psia) Bulk density Frozenb Shifting ◦F K

Ethylene oxide 300 0.887 166 175 1860 1288

Hydrazinec 1000 1.01 155 174 1077 854

Hydrogen peroxide (90%) 300 1.44 132 165 1370 1016

MHF-5c 1000 1.011 — 215 1916 1320

Nitromethane 300 1.13 218 225 3950 2450

n-Propyl nitrate 300 0.935 — 179 1840 1278

Otto Fuel II 1000 1.23 — 208 2042 1390

a Courtesy of Rocketdyne, Division of Rockwell International, and the U.S. Air Force Rocket Propulsion Laboratory.b Expanded to sea level, 14.7 psia (101.4 kN/m2).c Ammonia dissociation of 40% assumed.

−228.6◦ F (128.4 K), is sufficiently high so with properenvironmental control in appropriate spacecraft OF2

can be maintained in the liquid state with no boil-off,excessive pressure buildup, or need for refrigeration.Experimental rocket engine firings at the 100–400 lbf(445–1779 N) level, using diborane as the fuel wereconcluded successfully. Oxygen difluoride delivers about2% less performance than fluorine with hydrogen asthe fuel, and a little higher with a hydrocarbon fuel.It is a powerful oxidizing agent like fluorine and theinterhalogens, and the same care in material selection,cleaning, and passivation should be exercised. There issome indication that OF2 is less aggressive than fluorinewith nonmetallic materials. Thus, OF2 might be a betterchoice for some applications requiring gaskets, dynamicpump seals, and so on, since there are no nonmetallicsrecommended for use with flowing liquefied fluorine.

Other oxygen fluorides, O2F2 and O3F2, are thermallyunstable, decomposing below −71◦F (216 K).

4. Nitrogen Trifluoride

NF3 is the first of the N–F family of oxidizers. It also is acryogen and has lower performance than that of fluorine.Since NF3 is less reactive than fluorine, it was selected asan oxidizer for gas dynamic laser development. A widerange of metals and plastic materials are suitable for ni-trogen trifluoride service. With most fuels, NF3 is not hy-pergolic and offers only a marginal performance gain overthe much cheaper liquefied oxygen.

5. Oxygen

Since Professor Robert H. Goddard pioneered the use ofliquefied oxygen (O2) as a rocket oxidizer in 1926, ithas been the “work horse” oxidant for missiles and spacelaunch vehicles. Called A-Stoff, it was used in the German

V-2 missile and later in the U.S. Thor, Jupiter, Titan I, andSaturn. It is currently used in the Atlas, launch vehicleDelta, and Space Shuttle. Liquefied oxygen is nontoxic,stable, noncorrosive, inexpensive, and is readily availablefrom commercial sources. Being a cryogen, it is not desir-able for missile applications and extended space missions.

B. Storable Oxidizers

Because of the inherent difficulty of maintaining cryo-genic oxidizers in a state of readiness, the military has hada particular interest in storable oxidizers. Although thereis no specific definition of “storable,” it is usually consid-ered to encompass the ability of the oxidizer to remain ina ready condition within the missile or satellite for a min-imum of 5 years, and in some cases longer than 10 years.During the 1960s the military services were interested intactical weapons that required propellants with a low vaporpressure, <100 psig (689.5 kN/m2) at 160◦ F (344 K); alow freezing point, –20◦F (244 K) to –65◦F (219 K); benoncorrosive; and would not decompose. Less severe re-quirements are placed on silo-based strategic weapons andspace vehicles.

1. Nitrogen Fluorides

Tetrafluorohydrazine, N2F4, is also called dinitrogentetrafluoride. While this oxidizer can be kept liquid un-der its own vapor pressure, the pressure is high, 400 psiaat 70◦F (2.76 MN/m2 at 294 K), thus N2F4 is often con-sidered space storable rather than earth storable.

Tetrafluorohydrazine first became of interest to therocket community in 1957 as an intermediate leadingto a large family of postulated N–F compounds thatwould make improved storable oxidizers. These hypo-thetical materials had attractive theoretical performance,and the estimated physical properties appeared desirable.

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Under Project Principia, funded by the Advanced Re-search Projects Agency (ARPA), and programs supportedby the military services, a large synthesis effort was ini-tiated in the quest for high-energy, storable propellants.Among those successfully synthesized and characterizedwere C(NF2)4, FC(NF3)3, HNF2, ONF, ONF3, O2NF, andFNO3.

While many of the properties of these materials metexpectations, the family as a whole was found to be verysensitive and the compounds readily decomposed, oftenviolently. Extensive stabilization efforts were not fruitful;however, this research has led to a number of advancesin solid propellant chemistry. N2F4 has proved to be avaluable intermediate for the synthesis of solid oxidizersand energetic binders for rocket motors.

Mixtures of N2F4 with other oxidizers were also inves-tigated in an attempt to reduce its vapor pressure and yetretain its relatively high performance. N2F4 is a strongfluorinating agent resulting in unwanted reactions withmost additives and oxidizers. Some of the mixtures wereunstable and none proved to be useful. As a liquid oxi-dizer, the disadvantage of tetrafluorohydrazine for tacticalweapons is the rapid lowering of density as its low criti-cal temperature of 96.8◦F (309 K) is approached (Fig. 3),complicating rocket engine design. Neat N2F4 remains aviable liquid oxidizer candidate for space missions wherethe upper temperature limit is moderate. N2F4 is not reli-ably hypergolic with common rocket fuels.

2. Interhalogens

This family represents the highest energy earth-storableliquid oxidizers. They are toxic and highly reactive. Yet, inproperly prepared containers these materials exhibit out-standing storability characteristics.

a. Chlorine Pentafluoride. Chlorine pentafluoride(ClF5) is a product that resulted from the intensive fluo-rine chemistry research programs. It was first synthesizedin 1962, and fired in a small rocket thruster in 1964. It hasexcellent density and the best performance of the availablestorable oxidizers, which makes it attractive for volume-limited systems. Because of its reactivity, it is best suitedto prepackaged systems. The military has been reluctantto use ClF5 because of its unforgiving nature; an inadver-tent release of the material produces a toxicity hazard andwould likely result in a fire. ClF5 has been extensivelyevaluated in a variety of rocket engines and subsystems.Chlorine pentafluoride was the selected oxidizer for theAdvanced Liquid Axial Stage (ALAS) built by Aerojet fora Strategic Defense Initiative space-based interceptor. Thehigh performance and high density would serve to mini-mize both the weight and volume of systems that would

have to be orbited. It has been stored in mild steel contain-ers for more than 20 years with no signs of deterioration orcorrosion. The properties of chlorine pentafluoride mirrorthose of chlorine trifluoride, a product that has been pro-duced and shipped commercially for more than 40 years.ClF5 is not in current production; however, it is easilymanufactured from chlorine and fluorine. The more avail-able ClF3 is used as a substitute for ClF5 in test programswhen actual performance measurements are not needed.A complete ClF3 propulsion system was demonstrated fora liquid version of the Condor missile.

b. Perchloryl Fluoride. Perchloryl fluoride (ClO3F)is a stable derivative of perchloric acid first synthesized in1951. It is less toxic and much less reactive than ClF3 andClF5. ClO3F has the advantage of having some oxygenthat burns preferentially with carbon. Thus, ClO3F/MMH(monomethyl hydrazine, N2H3CH3) has an Isp close tothat of ClF3/hydrazine. The density of perchloryl fluorideis lower than that of ClF3, and because of its high coeffi-cient of expansion, the density drops rapidly at the highertemperatures of interest, 160◦F (344 K). It is not reliablyhypergolic with hydrazine and other fuels. With the emer-gence of ClF3, ClO3F no longer was of interest.

c. Oxychlorine Trifluoride. Oxychlorine trifluoride(ClF3O), called Florox, is an oxidizer that held promise.It boils at 87.1◦F (303.8 K), and has a density of 1.852.Because of its oxygen content, Florox is especially goodwith carbon-containing fuels (e.g., MMH). It is difficultand expensive to manufacture which has limited its avail-ability for rocket research programs.

d. Bromine Pentafluoride. Bromine pentafluoride(BrF5) has received attention as an oxidizer because ofits very high specific gravity (2.47). Nevertheless, its per-formance is mediocre.

3. Nitrogen Oxides

Nitrogen tetroxide (dinitrogen tetroxide) (N2O4) is the“work horse” storable oxidizer. Initially considered andevaluated for the Bomarc (but not selected for the pro-duction version), it found its first major application in theTitan II intercontinental ballistic missile. It is used in theTitan III, 34D, and IV space launch vehicles, the SpaceShuttle reaction control system, and in some of the newersatellite propulsion systems. The high freezing point ofN2O4 has restricted its use in many other applications.Produced by the fertilizer industry, N2O4 is available inlarge quantities at a low price.

Stress corrosion cracking of titanium tanks by N2O4 isinhibited by the addition of about 1 wt% nitric oxide (NO),

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the resulting formulation being called MON-1. To assurethat there will always be sufficient NO for protection sincesome NO is lost through venting, MON-3 (containing anominal 3 wt% NO) is used in the space shuttle and satel-lites.

NO depresses the freezing point of N2O4 without anIsp loss but with an appreciable increase in vapor pressure(Fig. 1). The freezing point of MON-25 is approximately–70◦F (216 K). Tetranitromethane will lower the freezingpoint to about –22◦F (243 K). Many attempts to find otherfreezing depressants have failed.

4. Hydrogen Peroxide

Hydrogen peroxide (H2O2) was of interest to Germany asbeing better suited for military purposes than liquid oxy-gen. A number of missiles and aircraft were developedbeginning as early as 1936. Although wartime Germanywas able to produce only 80% H2O2 (called T-Stoff), itwas not as corrosive as nitric acid and did not producetoxic fumes. It also had the advantage of being a mono-propellant that could be catalytically decomposed to pro-vide hot gases (see Section III.F.5). The Messerchmidt163A Komet interceptor, a rocket-powered aircraft, wasfirst flown in 1937 using hydrogen peroxide monopropel-lant. A later version, Me-163B, powered by a bipropellantengine used H2O2 as the oxidizer with C-Stoff, a hydrazinehydrate-based fuel.

After World War II, the United States produced 90%H2O2, and later a 98% grade. Peroxide readily found usein missile gas generators, research aircraft, and satelliteattitude control systems, and as the oxidant in aircraft su-perperformance propulsion systems.

The U.S. Navy sponsored the development of peroxideas an oxidant for underwater propulsion, and as an oxidantin superperformance rocket systems for jet fighters. TheU. S. Air Force used a hydrogen peroxide/JP-4 bipropel-lant propulsion system to give additional performance tomodified NF-104 aircraft used by the USAF AerospaceTest Pilot School.

While NASA and the Air Force made extensive useof hydrogen peroxide, the high freezing point of 11◦F(261.5 K) and the slow thermal decomposition have lim-ited its use for many applications. Some attempts weremade to lower the freezing point with additives such asammonium nitrite, but with limited success. With the se-lection of proper materials and cleaning and passivationprocedures, the decomposition rate is about 0.5% per year.The 98% peroxide has a lower rate of about 0.1% per year,but a higher freezing point, 31◦F (272.6 K). The oxygengas that results from the decomposition can be releasedthrough relief valves and porous plugs if the resulting pres-sure cannot be tolerated. A 90% peroxide system may besealed and stored for one year and a 98% H2O2 system for

five years if proper precautions are observed. Both 90%and 98% H2O2 have been extensively evaluated with avariety of fuels including the hydrazines, kerosine, andpentaborane in 5000 lb f (22 kN) engines. As can be seenfrom Tables V and VI, the specific impulse penalty is 25sec for H2O2/kerosine compared to LOX/kerosine and 10sec compared to N2O4/MMH. Recent thought is that thelower toxicity would make a lower cost system even withthe slow decomposition problem and the lower perfor-mance.

In the 1990s, NASA and others have experimented withH2O2/JP-8 engines of about 10,000 pounds thrust. Themotivation is to avoid the necessity of handling and storingcryogenic liquid oxygen. The downside is a loss of about25 sec of specific impulse which makes the vehicle heavier.

5. Fuming Nitric Acids

Various grades of nitric acid have been used extensively asstorable oxidizers for many years. They have a good liquidrange and are hypergolic with most amine-based fuels. Be-cause of their corrosiveness and tendency to thermally de-compose, considerable effort was expended by Germanyand the United States to overcome these traits. Water en-hances corrosion; therefore, early production centered onmaking 100% HNO3, known as white fuming nitric acid(WFNA), and keeping it absolutely dry. The Germans useda mixed acid, consisting of WFNA plus 10–17% sulfuricacid (a nitrating agent used in the explosives industry),which improved hypergolicity, and was believed (erro-neously) to be less corrosive. They also found that theaddition of 6% NO2 to WFNA, making red fuming nitricacid (RFNA), improved stability.

After exploring the decomposition mechanisms of fum-ing nitric acid, several investigators found that self-dissociation of nitric acid occurs, forming water and oxy-gen gas:

2HNO3 → 2NO2 + H2O + O2 (18)

Other studies to find freezing point depressants and meansto improve thermodynamic stability lead to the obviousconclusion that one could add NO2 and water to WFNAand force Eq. 18 to the left. An interesting developmentwas the finding that hydrofluoric acid (HF) reduced the ni-tric acid corrosion rate by a factor of 10 both in aluminumand stainless steel containers. From this work a family offuming nitric acid types was developed (Table X). TypeIIIA, containing a nominal 14% NO2, 2% water, and 0.6%HF, became the standard nitric acid oxidizer, and is com-monly called IRFNA (inhibited red fuming nitric acid).This composition was found to be the least corrosive andmost stable. It was used subsequently in the prepackagedBullpup and Lance missiles, the Bomarc, and in the Agenaspacecraft. Bullpups were still functional after 18 years

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TABLE X Fuming Nitric Acids—Chemical Composition Limits (Weight Percent)a

Nitrogen Nitric acid Solid as HydrogenType dioxide (NO2) Water (H2O) (HNO3) nitrates fluoride (HF)

IIIA 14 ± 1.0 1.5 to 2.5 81.6 to 84.8 0.10 max. 0.7 ± 0.1

IIIB 14 ± 1.0 1.5 to 2.5 81.6 to 84.8 0.04 max. 0.7 ± 0.1

IIILS 14 ± 1.0 0.5 max. 83.7 to 86.4 0.04 max. 0.7 ± 0.1

IVb 44 ± 2.0 0.5 max. 52.7 to 57.4 0.04 max. 0.7 ± 0.1

a From MIL-P-7254F Amendment 2.b Additional requirement for Type IV: Fe2O3 0.002 wt% maximum.

of unattended storage, and the Lance had more than 15years of successful deployment. Type I (WFNA), used inthe Rascal, and Type II (RFNA with 6% NO2 and no in-hibitor), used in rocket-assisted takeoff units (RATO), areno longer used and were dropped from the specification.

The maximum amount of NO2 that can be added tofuming nitric acid without a phase separation is 44%.This composition has the highest density, about a 3%specific impulse improvement over Type IIIA, and ismore corrosive. Officially identified as Type IV, is is alsocalled High Density Acid (HDA) and Maximum DensityIRFNA (MDIRFNA). It was used in some models of theAgena where the total operational period was measuredin months. Phosphoric pentafluoride (PF5) was found tobe a superior corrosion inhibitor over HF but was neveradopted for system use.

6. Tetranitromethane

Tetranitromethane (TNM), C(NO2)4, is a monopropellantthat has high specific gravity (1.368) and an oxidizing po-tential similar to that of N2O4, but it freezes at 57.4◦F(287.2 K). Some batches of TNM were sensitive to me-chanical shock, while others readily passed safety tests,indicating the possible presence of impurities. A eutecticmixture of 64% TNM and 36%, by weight, N2O4 freezesat –22◦F (243 K) and is less sensitive than neat TNM.

7. Other Oxidizers

Perchloric acid has received limited attention as a liquidoxidizer as perhaps being less corrosive than fuming nitricacid, yet having a low freezing point. At concentrationsapproaching 100%, HClO4 is very unstable and its perfor-mance is inferior to that of nitric acid. Mixtures of aminesalts of perchloric acid were studied briefly.

C. Cryogenic Fuels

1. Hydrogen

Liquefied hydrogen (H2) has high heat release and pro-vides very low molecular weight combustion products,

making it the fuel of choice with either oxygen or fluorinewhen its cryogenic nature is acceptable. Liquefied hydro-gen was used in the S-II and S-IVB stages of the SaturnV launch vehicle for the Apollo program.

Liquid hydrogen is operational in Centaur, the SpaceShuttle, and the upper stage of Ariane 5. Boeing is devel-oping the RS-68 engine for the first stage of the EELVwhich will use LOX/LH2.

Derived from petroleum sources, hydrogen is readilyavailable at a reasonable cost. Liquefied hydrogen existsas a mixture of ortho- and para-hydrogen. Ortho-hydrogenreverts exothermically to the para form, increasing boil-off. This problem is avoided by catalytic ortho to paraconversion by the manufacturer.

Slush hydrogen, a mixture of frozen and liquefied hy-drogen, and hydrogen subcooled below its boiling point,provides increased density and storability. This is espe-cially attractive for space vehicles to minimize structureweight and reduce the initial hydrogen boil-off rate. Thedensity values at the triple point of hydrogen are presentedin Table XI. The latent heat of fusion of solidified hydro-gen at the triple point is 25.02 Btu/lb (117.32 J/mole).Solid cryogenic storage can double the in-space storagelife of hydrogen.

2. Cryogenic Hydrocarbons

Liquefied methane (CH4), ethane (C2H6), ethene (C2H4),and propane (C3H8), both neat and in various mixtures

TABLE XI Density Values for Slush, Liquid, andSolid Hydrogen

State lb/ft3 kg/m3

At the normal boiling pointa

Liquid 4.425 70.889

At the triple pointb

Saturated liquid 4.81 77.0250% liquid/50% solid 5.09 81.54Solid state 5.41 86.64

a Normal boiling point: –423.3◦F (20.4 K) at 1 atm.b Triple point: –434.8◦F (13.8 K) at 1.021 psia (7.04

kN/m2).

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are excellent fuels, having better performance and com-bustion characteristics than kerosine, and the disadvantageof being cryogenic. The higher density of these liquefiedgases offer an advantage over hydrogen in very large ve-hicles by keeping the tank size down, and thus overalldry weight, although there is also a reduction in specificimpulse. Slush and supercooled states provide increaseddensity and longer storability.

D. Storable Fuels

1. Storable Hydrocarbons

Much attention and effort has been devoted to develophydrocarbon-based rocket fuels because of the low costand widespread distribution of petroleum products. Kero-sine is the most widely used fuel, generally with liquefiedoxygen as the oxidizer. This combination is used in theAtlas and Thor now launch vehicles, and was used in theTitan I and the Saturn V first stage. Propellant-grade kero-sine (called RP-1) is essentially JP-4 jet engine fuel witha controlled aromatic content to eliminate coke formationin regenerative cooled engine jackets. It also has a narrowdensity range needed for accurate volumetric loading ofmissile tanks.

Some hydrocarbons decompose under high tempera-ture. If the temperature is too high in the regenerativecooling passages such decomposition will leave carbonbehind to the extent that the cooling passages could be-come clogged. Even if not clogged a layer of carbon canreduce the heat transfer rate inside the cooling passagesand result in a burn through. Reduction of heat transfer ratewith a coating of carbon is is actually desired in some hy-drocarbon engines where a coating of carbon on the com-bustion chamber side of the cooling passage reduces theheat transfer rate. Where the engine is not regenerativelycooled by the fuel, it is much cheaper to use JP-8, a com-mon military jet fuel rather than the more expensive RP-1.

Dr. Goddard used motor-grade gasoline in his experi-ments. Decalin, diethylcyclohexane, octane, and aviationgasoline were also used as rocket fuels in the early yearsof rocketry in both the United States, and Germany. Dur-ing the 1950s the U.S. Air Force selected jet fuel to beused with fuming nitric acid in RATO units and super-performance propulsion systems to be used on the B-47,F-86, and other aircraft of that period. The JP-4/nitricacid system was plagued with combustion instability, hardstarts, and overall unreliability. Various additives, notablyamines and metal salts, were tried to reduce ignition delayand improve the combustion process, with limited suc-cess. The U.S. Navy had fared better with its JP-5/H2O2

engine development program. Subsequently, a hydrogenperoxide rocket system (using the onboard JP-4 fuel) was

incorporated into NF-104 aircraft used for astronaut train-ing at the Aerospace Test Pilot School, Edwards AFB,California.

Both the U.S. Air Force and Navy funded efforts todevelop petroleum-derived fuels for rocket engines usingnitric acid as the oxidizer. Many different compounds wereinvestigated including mercaptans and thiophosphates.The individual components in JP-4 were isolated and eval-uated on their ability to improve or to hinder hypergolicityand combustion. Hydrazine compounds as additives werefound to be the answer. Many other hydrocarbons (e.g.,aromatics, aliphatics, acetylenes, terpenes, ethers) weretried in many propulsion systems. Turpentine was used inthe French Diamant launch vehicle. Another fuel of inter-est was the German Viscol, a vinyl isobutyl ether. Mosthydrocarbon fuels offer about the same specific impulse.Thus, the choices are factors of availability, physical prop-erties, and cost. For large boosters and launch vehicles, ahigh propellant density is desirable to reduce overall size,and thus structure weight. Shelldyne

❤™ -H, hydrogenateddimer of norobornadiene, is an example of a dense fuel(1.11 kg/m3) originally developed as a thermally stable,endothermic fuel ramjets as RJ-5.

Several variations of kerosine have arisen in the 1990s:(1) RP-2 a mixture of kerosine or heptane with quadri-

cyclane; (2) RG-1 the Russian version of kerosine rocketfuel with higher density than RP-1; and (3) JP-8X andJP-8 + 100.

Other alternative hydrocarbons under research and de-velopment which have a 10–15 sec specific impulse advan-tage over RP-1 include bicyclopropylidene, spiro-pentaneand tripropargyl-amine. A 10-sec increase over LOX/RP-1is available by using LOX/UDMH as did the Juno rocketin 1957. However, the current preference is to search forsomething less toxic than the hydrazine family to enhancethe performance of an existing kerosine fueled vehicle.

RP-2 uses quadricyclane, which releases energy from itsmechanically strained molecule in addition to the energyof combustion. Quadricyclane is strained due to the cy-clopropane rings. Compared to LOX/RP-1, a 50% quadri-cyclane/50% RP-1 mixture, the specific impulse gain is3 sec. In a 65% quadricyclane/35% n-heptane, the specificimpulse gain is 5 sec. For 100% quadricyclane, the specificimpulse gain would be 7 sec but experiments have revealedrough combustion. Other strained ring compounds such ascubane are also being studied for performance improve-ment over RP-1 in existing vehicles

2. Alcohols

In Germany during the late 1930s and through World WarII, methanol (CH3OH) and ethanol (C2H5OH) (M-Stoffand B-Stoff) were widely used as rocket fuels, being more

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available than petroleum products. The V-2 used ethanoland liquefied oxygen. Captured German technology andequipment laid the foundation for alcohol-fueled rocketsin both the United States and the Soviet Union after thewar. The alcohol fuels have excellent storability and com-bustion properties, but relatively low performance. Fur-furyl alcohol (C4H3CH2OH) improves the hypergolicityof hydrocarbon and amine fuels with nitric acid.

3. Amines

The German emphases on storability and hypergolicityled them to the amines. Tonka-250 (R-Stoff, used in theTaifun Missile) consisted of 57% xylidine and 43% tri-ethylamine. Tonka-500 contained a number of compo-nents: toluidine, triethylamine, aniline, gasoline, and ben-zene. C-Stoff, used in the Me163B Komet consisted ofmethanol, 57%; hydrazine hydrate, 30% (reduced to 15%at the end of WW II to stretch supplies); water, 13%; andcopper, 30 mg/l (as potassium cuprocyanide to improvecombustion).

The U.S. Army had similar interests, plus a desiring fora low freezing point. To this end, the Jet Propulsion Lab-oratory investigated a large number of additives to makegasoline hypergolic. At least 50% of xylidine or aminewas needed to impart hypergolicity. Furfuryl alcohol wasfound to be a good fuel component. The final result wasMIL-P-45700, a mixture of 51% aniline and 46.5% fur-furyl alcohol that froze at –45◦F (230.4 K). This mixturewas used in the Navy’s Lark and in the Army’s Corpo-ral missiles. Later, 7% hydrazine was added to improvestorability and low-temperature hypergolicity.

A bewildering list of amines and mixtures has been pre-pared as rocket fuels and evaluated by various organiza-tions. A commercial grade of monoethyl aniline (N -ethylaniline) exhibited a freezing point of −81◦F (210 K) andexcellent hypergolicity with fuming nitric acid. Diethy-lene triamine was of interest as a density additive. Acetoni-trile improved the low-temperature viscosity of amines.None of the possible hydrocarbon/amine fuels had a per-formance much better than the others; thus, the applicationof these fuels was often based on availability, hypergolic-ity, freezing point, and smoothness of combustion.

4. Hydrazines

Captured stocks of World War II hydrazine hydrate openeda new era of propellant research in the United States. Thetechnology to manufacture anhydrous hydrazine was de-veloped, a capability not available in wartime Germany.The high freezing point of anhydrous hydrazine and its ten-dency for chamber explosions have restricted the use of theneat material to small space engines, where it is the mono-

propellant of choice. Unsymmetrical dimethylhydrazine(UDMH) and monomethyl hydrazine (MMH), N2H3CH3,became the focus of considerable interest in the 1950s andare now the principal fuels used in storable liquid rocketpropulsion systems. Wide acceptance of UDMH was slowbecause of the interest from the U.S. Navy and Air Force touse petroleum-derived fuels. Initially UDMH was used asan additive (17% UDMH) to JP-4 to stabilize combustionin the Nike Ajax. This mixture was called JP-X Type II.A 40% UDMH/60% JP-4 blend, JP-X Type I, was used inthe Bomarc. JP-X is covered by MIL-P-26694.

A series of mixed amine fuels (MAF) were developedfor tactical missiles and launch vehicles. (See Table XII.)MAF-1 was used in the Bullpup deployed by the U. S.Navy and Air Force. MAF-4, also known as U-DETAand Hydyne, was used in the Jupiter-C and target drones.A family of mixed hydrazine fuels (MHF) was also de-veloped, primarily as monopropellants but also used asbipropellant fuels (Table XIII). MHF-3 was developed asthe fuel for the Condor, and was used as M-86 monofuelin aircraft auxiliary power units (APUs). MHF-5 (laterMHF-5B) was selected for the Condor gas generator. BA-1014 and BA-1185 were tailored for fluorine or interhalo-gen oxidizers. The water in these fuels was added to pro-vide oxygen to combust preferentially with the carbon inMMH, eliminating free carbon in the exhaust.

A 50–50 wt% blend of anhydrous hydrazine andUDMH (often called Aerozine-50) is used in the Titan II,III, IV, and derivatives. A low freezing point is provided,and the stability of hydrazine is enhanced. The 50–50 mixwas chosen over MMH because of lower cost and bet-ter availability. The cost difference no longer exists, andthe production of MMH matches current demand. There-fore, MMH is now the preferred storable fuel for newpropulsion systems. Neat UDMH was used in the Agenaprogram and in the U. S. Army’s Lance. Now MMH isused in the space shuttle reaction control system engines,the Minuteman, and Peacekeeper postboost propulsionsystems, and is being selected for satellite bipropellantpropulsion systems. The hydrazines have exhibited excel-lent storability, with no degradation for more than a 20year period.

Symmetrical dimethyl hydrazine freezes at 16◦F(264.2 K) and has no advantages over UDMH. Ethy-lene dihydrazine (EDH) offers a higher density, makingmixtures of EDH and hydrazine with N2O4 as the oxi-dizer attractive candidates for launch vehicle upper stages.Aerozine-67 (67.2% EDH/32.8% hydrazine, by wt.) pro-vides a substantial payload increase over that of neat MMHor Aerozine-50, although the Isp is slightly lower.

Ammonia is an excellent fuel, but its high vaporpressure has restricted its use. In 1949 JPL first eval-uated ammonia with fuming nitric acid and N2O4. An

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TABLE XII Mixed Amine Fuel Compositions (Nominal)a

Fuel DETAb Acetonitrile UDMH PRODGAc DIPGAd TIPRAMe Water f

MAF-1 50.5 9.0 40.5 — — — —

MAF-2 — — — 42.6 48.3 9.1 —

MAF-3 80.0 — 20.0 — — — 1.0

MAF-4 (U-DETA) 40.0 — 60.0 — — — —

MAF-5 50.5 20.0 29.5 — — — —

a By weight percent.b Diethylenetriamine.c Propargyl diglycidylamine.d Dipropargyl glycidylamine.e Tripropargylamine.f Propellant-grade DETA and UDMH normally contain water (<1.0% by weight), thus all mixtures will have a small

amount.

ammonia/liquefied oxygen engine powered the X-15 re-search aircraft in the XLR-99 engine.

5. Metal-Based Fuels

The known liquid metal alkyls are low in performance, butthey are pyrophoric. Triethyl aluminum and triethyl boronare useful as igniters in liquefied oxygen/kerosine propul-sion systems (e.g., Atlas). Of the light metal hydrides,only diborane and pentaborane are normally liquid. Dib-orane is especially attractive for use with Flox or OF2 inspace-storable propulsion applications. B2H6 decomposesabove −4◦F (253 K), necessitating the use of refrigerationfor extended storage periods. Pentaborane is very stableas demonstrated by desert storage in mild steel cylinders

TABLE XIII Mixed Hydrazine Fuel Compositions (Nominal)

Freezing Hydrazine MMH Hydrazinium Water a

Fuel Point, ◦F(K) (wt%) (wt%) nitrate (wt%) (wt%)

MHF-1 −65(219) 23.3 43.3 31.4 —

MHF-2 −65(219) 23.3 36.5 40.2 —

MHF-3 (M-86) −65(219) 14.0 86.0 — —

MHF-4 −43.6(231.2) 32.5 50.5 17.0 —

MHF-5 −44.6(230.6) 26.0 55.0 19.0 2.0 max.

MHF-5A <−65(219) 13.0 68.0 19.0 —

MHF-5B −45(230.4) 23.0 58.0 19.0 —

MHF-6 <−65(219) 16.5 73.0 — 10.5

MHF-7 −65(219) 14.0 81.0 — 5.0

MGGP-1 −65(219) 63.0 — 10.0 27.0

H-70 (MGGP-3) −58(223) 70.0 30.0 — —

TSF-1 <−65(219) 60.0 — 21.0 19.0

TSF-2 <−65(219) 58.0 — 25.0 17.0

BA-1014 −2.5(254) 66.7 24.0 — 9.3

BA-1185 −65(219) 29.8 50.5 — 19.7

a Propellant-grade hydrazine and MMH normally contain water (<1.0 percent by weight),thus all mixtures will have a small amount.

for more than 30 years. Complete combustion with lique-fied oxygen and hydrogen peroxide is difficult, resultingin slow burning condensed boron particles and the for-mation of vitreous boron oxide solids in nozzles. Sinceboron fluoride is gaseous, these problems are eliminatedwith fluorine-based oxidizers. A unique propellant com-bination, pentaborane/hydrazine was also studied. In thiscase, the hydrazine acts as the oxidizer. The high theoreti-cal Isp of this system was predicated upon the formation ofboron nitride (BN). Combustion inefficiencies caused bycondensed products, kinetic limitations, and combustioninstability made the realization of high delivered perfor-mance difficult.

Experimental fuels based on aluminum and lithiumborohydride were evaluated, but were dropped when

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UDMH became available. Beryllium borohydride hasbeen tested, but its toxicity precluded further consider-ation.

The hybalines are a group of coordination compoundsof light metal hydrides or metal borohydrides, complexedwith Lewis bases such as amines or ethers. Following aresome of the compositions evaluated:

Hybaline A-3: methylamine aluminum borohydrideHybaline A-4: dimethylamine aluminum borohydrideHybaline A-5: 53 wt% A-3 and 47 wt% A-4Hybaline A–14: 2-ethylhexylamine aluminum

borohydride

Hybaline B fuels, amine adducts of beryllium borohy-dride, were also prepared and studied. Hybalines may beviewed as a densified form of hydrogen bounded by en-ergetic light metal atoms. This family of fuels are hyper-golic with most rocket oxidizers. They react with water,and sometimes are pyrophoric in air. Propellant feed sys-tems must be kept very clean since traces of moisture andair would form solid materials, causing flow problems.Hybalines are neither sufficiently stable nor storable formilitary applications.

Lithium has attractive theoretical performance withfluorine-based oxidizers. Experiments have been con-ducted using molten lithium injected into a 3000 lb f

(13.3 kN) thrust chamber with an interhalogen oxidizer.Solid lithium was used in hybrid experiments with chlo-rine trifluoride as the oxidizer. In this case, heat transferto the solid metal must be controlled to prevent prematuremelting, yet sufficient vaporization of the fuel must occurto sustain proper combustion. Lithium alkylates, for exam-ple, butyllithium, contain insufficient lithium to enhanceperformance.

E. Heterogeneous and Gelled Propellants

Gelation of liquid rocket propellants is of interest for sev-eral reasons:

(1) retardation of evaporation rate of volatile or toxicliquids; (2) retardation of mixing of spilled propellantsto prevent explosions; (3) suspension of energetic solidpowders to improve performance; and (4) reduction ofsloshing in tanks.

The above advantages come with some unique prob-lems:

1. Difficult control of mixture ratio because of theheterogeneous flow properties

2. Slightly decreased delivered performance comparedto theoretical

3. Higher pressure drops causing heavier inert weight

4. Decreased service life at temperature extremes5. Increased unusable residual propellants further

penalizing mass fraction.

Light metals have high heat release, making them at-tractive as fuels or fuel components. As mentioned previ-ously, there are very few candidates that are liquids. Oneway to use high energy, solid materials is to suspend themin a liquid carrier. In particular, Al, AlH3, B, Li, LiH, Be,and BeH2 have been of most interest. Considerable re-search was undertaken to find techniques that would keepthe solid from settling. Very fine particle size (typically0.1–10 µm) is needed and the liquid thickened. Differenthydrocarbon, hydrazine, and alcohol carriers were evalu-ated. Hydrazine and MMH were of most interest becauseof their higher energy, but were subject to gassing exac-erbated by the high surface area of the metal or hydridepowder. Gels, slurries, and emulsions were formulated andevaluated.

Thixotropic gels were of particular interest because oftheir ability to be liquefied by internal mechanical stressand to return to the gel state when the stress is removed.The consistency of thixotropic gels depends on the du-ration and rate of shear. Under static conditions, thesegels exhibit a weak structure and are nonflowing, semi-solid substances. If, however, the yield point is exceededand the gels are sheared at a constant rate, the structurewill progressively weaken and the viscosity will decreasewith time. The structure of true thixotropic materials willbreak down completely under the influence of high shear-ing action such as injection through an orifice, in whichcase they behave like true Newtonian liquids. As soonas the shearing action has ceased, however, the struc-ture will reform. The rate of breakdown and reformationof the structure of a gel is a characteristic of that gel,and the time required for the transition may vary froma fraction of a second to many hours. Many gelling andthickening agents were tried, with colloidal silicon diox-ide, hydroxyethyl cellulose, and polyacrylic acid beingtypical. Lithium and its compounds proved too reactivefor a stable propellant. The toxicity of beryllium and itshydride severely restricts its usefulness. Boron gels arebest suited for air-breathing engines, such as the ductedrocket. Gels containing AlH3 evolved too much hydrogengas to be practical. Several successful aluminum-basedformulations are given in Table XIV. Alumizine, coveredby MIL-P-27412 (subsequently cancelled) was fired ina modified Titan II engine. MICOM GEL was tested inan Lance engine and several organizations have evalu-ated NOTSGEL-A in a variety of thrust chambers withdifferent oxidizers. These formulations meet long-termstability requirements and present no undue operationalhazards.

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TABLE XIV Heterogeneous Rocket Fuel Composition

Component Alumizine-43G NOTSGEL-A MICOM GEL

Aluminum Powder 43.0 60.0 60.0

MMH — 31.1 38.5

Hydrazine 56.3 5.1 —

Colloidal Silica — 3.0 —

Modified Polyacrylic Acid 0.3 — —

Water 0.4 — —

Hydroxypropyl Cellulose — — 1.4

Hydroxyethyl Cellulose — 0.8 —

Dimethylurea — — 0.1

The gelling of cyrogenic propellants has also been stud-ied. Liquified hydrogen can be gelled with light hydrocar-bons (e.g. methane, ethane), which are solid at hydrogentemperatures.

Gelation of liquid hydrogen has a payoff of reducingboil-off rate and significantly reducing spill radius. A 10%density increase is possible with gelled hydrogen by theaddition of 10% solid ethane or methane powder.

Current research and development activities includegelation of RP-1 in order to add aluminum powder to aLOX/RP-1 system to gain increased performance.

1. Oxidizers

Liquid rocket oxidizers are more energetic than are solidoxidizers; thus, there are no solid components that canbe added to a liquid oxidizer to enhance specific impulse.However, potential for improvement in safety and opera-tional usage is the incentive to develop gelled oxidizers. Inan accidental spill, gelled oxidizers would have reducedflow characteristics and lower vaporization rates, makingit easier to contain the spilled propellant and rendering iteasier to clean up. Alleviation of sloshing in tanks providesgreater flight stability of missiles.

Another reason to gel oxidizers is to match the flowproperties of a gelled fuel. The addition of LiNO3 to nitricacid gelled with silica, for example, increases the density,does not seriously detract from Isp, and provides rheo-logical characteristics similar to those of MICOM GEL,minimizing mixture ratio shift as propellant temperaturesare varied across the operational range.

Liquefied oxygen, oxygen difluoride, fuming nitricacid, and nitrogen dioxide have been gelled. Gelation ofLOX is actually possible using frozen N2O4 particles.

While the resulting formulations were chemically sta-ble and found insensitive to shock, long-term storabilityhas not been fully resolved. Syneresis, or the exuding ofliquid, is a problem that may occur after several monthsor a few years of storage. Rheological behavior can also

change with time. The following are typical gelled oxi-dizer compositions:

1. Oxygen and 5 wt% aluma or titania2. Oxygen difluoride and 4.5 wt% colloidal silica3. RFNA and 4 wt% colloidal silica4. IRFNA and 5 wt% carbon black5. Nitrogen dioxide and 3.5 wt% colloidal silica6. MON-15 and 3.5 wt% colloidal silica

The gelling of chlorine pentafluoride (ClF5) with sub-micron particle size ferric oxide and alumna was not suc-cessful. Silica reacts with CLF5. Fluoride salts (e.g., CaF2,LiF) are chemically stable with ClF5, but large amountsare required for gelling.

2. Safety

Gelling of liquid propellants has been proposed as a meansof reducing the consequences of an accident in which pro-pellants are released from their container. A large-scalespill of liquid propellants will result in splashing andspreading of the liquid over a large area. Tests have shownthat under similar conditions the gelled propellant will re-main confined to a smaller area due to its high viscosity,making it much easier to clean up. In the case of a si-multaneous spill of a hypergolic propellant combination,the propellants will ignite at the interface, but will not beinvolved in a violent fuel/oxidizer deflagration. The fuelwill continue to burn as a normal fuel/air reaction, but be-cause of its immobility, will not spread. Increased safetyis also provided by gels when small, low pressure leaksoccur at fittings or cracks since gels will not flow from theleak as would a liquid.

Tests have also shown that gelation of liquids will alsoreduce the rate of vaporization, thus the spread of toxic orflammable vapors is lessened. However, the vapor pres-sure, reactivity, and toxicity of the propellants are not al-tered. In other tests, tanks containing gelled propellants

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subjected to bullet impact or external fire experiencedpressure rupture but did not detonate as will typical highperformance solid propellants.

F. Monopropellants

There are three basic types of liquid monopropellants, ormonofuels:

1. Single substances (molecular compounds) thatcontain an oxidizing agent and combustible matter inthe same molecule

2. Intimate mixtures of compounds that are stable undernormal conditions

3. Thermodynamically unstable chemicals

Upon ignition, the first two types of monopropel-lants will undergo combustion much like a bipropellantcombination. The components in the second type must becompatible at normal temperatures, but react when ignitedor heated under pressure to give hot combustion gases. Thethird type decomposes to produce hot gases by means ofa catalyst or a thermal source. These are the monopropel-lants, specifically hydrogen peroxide and hydrazine, thathave been the most acceptable and useful in a wide varietyof systems.

The challenge in developing high-energy monopropel-lants is finding compositions that are easy to handle andstable to a wide variety of adverse environmental influ-ences, yet will decompose rapidly and smoothly wheninitiated to deliver its energy in a rocket or gas generator.

1. Molecular Monopropellants (Combustion Type)

Examples of molecular liquid monopropellants that un-dergo a combustion process are nitromethane (CH3

NO2), ethylene oxide (C2H4O), n-propyl nitrate (C3H7

NO3), ethyl nitrate (C2H5NO3), and tetranitromethane[C(NO2)4]. Electrical sparks and pyrotechnic squibs areused to ignite these propellants.

Ethylene oxide, a commercial chemical, was used topower decoys. It is stable under high-pressure flow con-ditions; however, the vapor phase is sensitive. This com-pound is subject to exothermic reactions when in contactwith various acids, bases, salts, and catalysts. Decompo-sition of ethylene oxide is difficult to sustain because ofits high autoignition temperature of 804◦F (702 K) andcarbon formation from thermal cracking.

The U. S. Navy is developing cyclodextrin nitrate withan Isp of 240–280 sec and a density near 2.0 gm/cc.Dinitramine is another monopropellant candidate with anIsp of 270–300 sec and a density of 1.9 gm/cc. Anothermonopropellant candidate in the 1990s is hydroxyl am-monium nitrate.

2. Nitroaliphatic and Nitrate Esters

These compounds have had limited application as mono-propellants. Nitromethane, used by race car and modelairplane enthusiasts, has the highest performance and wasthe most studied. It is normally sensitive to shock and ther-mal stimuli, and has a high flame temperature of 3950◦F(2450 K). Engine tests were characterized by instabilityand inefficient operation. While most additives generallyincrease its sensitivity, a few alcohols and hydrocarbonswere found to improve stability and lower the flame tem-perature with a modest decrease of performance. The Ger-man Myrol, consisted of 80% methyl nitrate and 20%methanol, was reasonably safe to handle.

N -propyl nitrate (NPN) was used in jet engine startersof U.S. military aircraft and for decoys. The British usedthe similar isopropyl nitrate for engine starters and missilegas generators. NPN was found to be sensitive to adiabaticcompression. The addition of 60% ethyl nitrate resultedin an 8% specific impulse improvement, but further sen-sitizes NPN.

3. Composite Monopropellants

a. Amine-nitric acid. A major effort in the develop-ment of high-energy monopropellant candidates was di-rected at the amine nitrate-nitric acid composite-type for-mulations. These mixtures can be considered premixedbipropellants. This family consists principally of sec-ondary, tertiary, ditertiary, and quaternary amines that areconverted to the corresponding nitric acid salts and mixedin solution with nitric acid (WFNA or RFNA). A large va-riety of mixtures has been made and evaluated, sometimeswith disastrous results. Examples of this class that havebeen studied are

1. WFNA and pyridinium nitrate (Penelope)2. WFNA and diisopropylaminium nitrate (Isolde)3. WFNA and [1,2 4-trimethyl-1,4-diaza bicyclo

(2,2,2)] octane dinitrate (Cavea)

This family of composite monopropellants has moder-ate theoretical performance levels ranging from ∼220 to260 sec (2158 to 2550 N sec/kg). The performance, sensi-tivity and thermal stability of these composites vary withthe amine structure and oxidation ratio. As the oxidationratio approaches stoichiometric, the performance and thesensitivity are increased and the thermal stability is de-creased. As a result, these composites were found to bequite shock sensitive and of questionable thermal stability.Cavea underwent extensive characterization and engineevaluation at the 10,000 lb f (44.5 kN) thrust level. Afterseveral explosions occurred, testing was discontinued.

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b. Difluoroamines. Most of the difluoroamino com-pounds and composite monopropelants have exhibitedpoor thermal stability and extreme shock sensitivity.The most extensively studied of this family was 1,2-bis(difluoroamino)ethane as both a molecular monopro-pellant and in solution with either tetranitromethane ornitrogen tetroxide (the latter mixture known as Cyclops).The mixtures were less sensitive than the neat compound,but not sufficiently to avoid detonations.

The BN-type monopropellants are perhaps the high-est energy (Isp about 300 sec) (2942 N sec/kg) consistingof borane compounds dissolved in hydrazine. Dekazine(54% diamino-decaborane and 46% hydrazine) receivedthe most attention, being fired in a 600 lb f (2669 N) thrustchamber at 1000 psia (6:895 MN/m2). Also in this familyare the Hepcats, [(CH3)3 NNH2]+B10H13-type salts dis-solved in hydrazine. Low-delivered performance and highsensitivity contributed to the demise of this family.

c. Otto fuel. The Otto monopropellants, a series offormulations based on desensitized aliphatic dinitroxylesters, were originally developed for torpedo propulsion.Otto fuel II is the most successful composition, havingoutstanding safety characteristics while meeting strin-gent temperature, shock stability, and energy require-ments. Under normal conditions, it is considered to benonflammable and nonexplosive, making it a safe pro-pellant. Its major constituent is propylene glycol dini-trate stabilized with dibutylsebacate (see Table XV). OttoFuel II is used in the Mk 46 Mod 1 and Mk 48 torpe-dos, and it has been selected for several developmentalprograms, including a liquid-propellant gun and a rocket-propelled target drone. Ignition is accomplished with asolid squib that pressurizes the chamber to the 200 psi(1.38 MN/m3) necessary to sustain decomposition of theOtto Fuel.

4. Heterogeneous Monopropellants

Another group of composite-type monopropellants arethe heterogeneous monopropellants that consist of high-energy or high-density solid ingredients suspended invarious liquid vehicles as a slurry or through gelation.

TABLE XV Composition of Otto Fuel II

Component Amount (wt%)

Propylene glycol dinitratea 75.8–76.2

2-Nitrodiphenylamine 1.4–1.6

Di-n-butyl sebacate 22.2–22.8

Sodium ion 0.8 ppm (max.)

a 1.2 Dinitroxypropane.

Essentially, there are three chemical types of heteroge-neous monopropellants that have been studied: Arco-gels, B4C/oxidizer gels, and beryllium composites. TheArcogels are thick slurries of aluminum powder andnitrate or perchlorate salts in an uncured solid binderthat do not shear thin. A typical formulation, APG-25,[63.2% NH4ClO4 + 20.7% Al + 14.8% (CH2)x + 1.3%thickener], has a density of 1.8 kg/m3 and an Isp of 166 sec(1.628 kN sec/kg). Its very high viscosity greatly compli-cated a feed and engine system design.

Heterogeneous formulations of B4C in various oxidiz-ers, including nitrogen dioxide, perchloryl fluoride, andnitric acid, gelled with carbon black or SiO2 were evalu-ated in small engines for prepackaged and torpedo appli-cations. Moderate performance, about 225 sec (2,207 Nsec/kg), was realized. These compositions are insensitiveto shock, but appear to be unstable at 160◦F (344 K).

The beryllium monopropellants consist of Be and BeH2

in solutions of hydrazine, hydrazinium nitrate, water, orhydrogen peroxide gelled with colloidal silica, aluminumoctanoate, or Kelzan-AO. One patented formulation hav-ing a theoretical performance of 310 sec (3040 N sec/kg)is:

15.32 wt% Be+53.91wt% N2H5NO3 +30.77 wt% N2H4

These compositions are insensitive to shock, but are sub-ject to poor long-term storability and gassing. Because oftheir toxicity and inefficient combustion, beryllium-basedpropellants are of little current interest.

5. Molecular Monopropellants(Decomposition Type)

a. Hydrogen peroxide. Hydrogen peroxide was thefirst widely used monopropellant. During the late 1930sand through World War II, the Germans used 70–80%H2O2, and called it T-Stoff. They used it to drive the bipro-pellant pumps in the V-2. After the war, the technology toproduce 90–99% H2O2 was developed. Hydrogen perox-ide was also used to drive pumps in the early U.S. missilesand in satellite reaction control systems. In the X-15 high-altitude research aircraft, peroxide was used to drive themain propellant pumps and APU, and in the attitude con-trol thrusters.

Hydrogen peroxide is decomposed catalytically to formsuperheated steam and gaseous oxygen:

100% H2O2 → H2O (steam) + 12 O2

+ 685 Btu/lb (54.10 kJ/mole)

or

90% H2O2 → H2O (steam) + 12 O2

+ 505 Btu/lb (37.96 kJ/mole)

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Many materials act as catalysts to decompose peroxide.The following have been successfully used in operationalsystems:

1. Calcium, potassium, and sodium permanganate in anaqueous solution (German Z-Stoff)

2. Permanganate salts deposited on ceramic or metalballs

3. Silver and other noble metals4. Transition metal oxides

Screens of silver or silver-plated brass are the com-monly used H2O2 catalysts in the United States.

In the search for higher energy and low-freezing mono-propellants, experimental compositions of hydrogen per-oxide containing various amounts of ethanol, ammonia,and ammonium hydroxide, singly and in combination,were tested but with adverse effect on sensitivity andstability.

b. Hydrazine. Hydrazine’s ease of decompositionand freedom from the instabilities of the other monopro-pellants has led to its becoming the principal fuel for low-thrust satellite propulsion. It has replaced hydrogen per-oxide in almost all monopropellant applications becauseof its better storability and higher performance. IntelsatIII and IV, Pioneer 10 and 11, Mariner 69, Landsat, FLT-SATCOM, Telestar, and Tracking and Data Relay Satelliteare a few of the many systems equipped with hydrazinethrusters. Hydrazine is also used to drive the APUs inthe space shuttle. H-70 (70% N2H4 and 30% H2O) pow-ers the emergency power unit (EPU) in the F-16 and F-20fighter aircraft.

Hydrazine is decomposed by either thermal or catalyticmeans, producing hot gases of ammonia, hydrogen, andnitrogen:

3 N2H4 → 4 NH3 + N2 +144,300 Btu/lb (10.74 MJ/mole)

The resultant ammonia product is also dissociated by hy-drazine catalysts. Since its decomposition is an endother-mic process, it must be minimized to avoid serious perfor-mance loss:

4 NH3 → 2 N2 + 6H2 − 79,200 Btu/lb (5.90 MJ/mole)

Typically, about 40% of the ammonia is dissociated, beingcontrolled by catalyst bed design and residence time.

Hypergolic slugs were used to initiate hydrazine decom-position in the early systems. Later, catalysts composed oftransition metals and their oxides were used which pro-vided a limited restart and pulse-mode capability. Thesecatalysts were heated electrically to about 800◦F (700 K)to effect rapid hydrazine decomposition. The availabilityin 1964 of Shell 405 catalyst, which causes spontaneous

decomposition of hydrazine, greatly expanded the applica-tions of monopropellant hydrazine. It opened the way forthe use of hydrazine in systems requiring rapid-responsemultiple starts, especially the 0.2–25 lb f (0.9–111 N)thrusters used on satellites. Shell 405 catalyst consists ofa porous aluma substrate impregnated with chemicallydeposited iridium. The catalyst granules contain about33 wt% Ir and are in the range of 14–30 mesh. The surfacearea of the catalyst is typically 70 to 120 m2/g.

HYDRAZINE BLENDS. Various compounds havebeen evaluated as means of improving hydrazine mono-propellant performance and to depress its freezing pointwithout adverse effect on catalysts. The most successfulof the formulations are listed in Table XIII. MHF-3 wasused in auxiliary and emergency power units on many ex-perimental and prototype aircraft. Carbon resulting fromthe MMH decomposition can reduce the gas generatorlife for some applications. The TSF series, developed asmonofuels for aircraft turbine engine starters, do not havethis potential carbon problem. The addition of oxygen inthe form of hydrazinium nitrate (HN) helps to alleviatecarbon formation. The MHF-5 series are very stable attemperatures below 140◦F (333 K) in properly preparedcontainers. HN also provides a small Isp gain over neathydrazine. Shell 405 is the catalyst typically used withthese blends. It functions at temperatures as low as −65◦F(219 K).

Hydroxylamine, NH2OH, will depress the freezingpoint of hydrazine, but causes excess gas evolution.Methoxyamine (O-methylhydroxylamine, MOA) andmethoxyamine nitrate (MOAN) are also good depressantswith acceptable stability and offer a small performanceimprovement. Hydrazoid-X consisted of 86% MOA, 12%N2H4, and 2% H2O. AMICOM propellant, 68% N2H4,20% MOAN, and 12% H2O, has a freezing point of −58◦F(223 K).

Binary, ternary, and quaternary mixtures using the fol-lowing compounds have been also evaluated: ammonia,water, MMH, hydrazine azide, and monomethylaminenitrate.

G. Tripropellants

The tripropellant concept uses the high heat release fromthe oxidation of a light metal or metal hydride to heathydrogen expanded to generate a very high specific im-pulse (see Table VIII). The hydrogen serves as an expand-able, inert working fluid of low molecular weight, not un-like a thermodynamic nuclear rocket. The hydrogen canbe added with the reactants, or if the hydrogen preferen-tially reacts with the oxidizer, it can be introduced after thecombustion zone. Hydrogen can be introduced as a metalhydride, making the system appear to be a bipropellant

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system; the differentiation being there is excess hydrogen(from a thermochemical standpoint) for expansion with-out being oxidized. Pentaborane + hydrogen peroxide isan example of such a system.

The chemical propellant systems with the highest theo-retical specific impulse are beryllium/ozone/hydrogen andlithium/fluorine/hydrogen. The Be/O2/H2 system was ex-perimentally evaluated with the liquefied oxygen gelledto suspend the powdered beryllium. Fluidized injection ofBe powder was also attempted. The technical difficultiesof incorporating the light metals and their hydrides into aliquid system has limited the development of a practicaltripropellant propulsion system.

H. Hybrids

A hybrid system uses a liquid propellant (usually the ox-idizer) with a solid propellant grain in the same rocketengine. The configuration of the solid fuel is similar tothat of a solid propellant motor. Many different propel-lant combinations can be used, ranging from cryogenicoxidizers to conventional storable oxidizers, with metal-loaded fuels as well as common fuels such as rubbers orplastics. As a rule, liquid oxidizers are superior to solidin specific impulse, while solid fuels, particularly metal-ized, offer a performance and density improvement overliquids. Because it is possible to select dense oxidizers,such as ClF5 with a specific gravity of 1.78 and a metal-loaded fuel with a specific gravity of 2.0, hybrid systemsappear attractive for volume limited applications. Hybridshave the flexibility of being restarted and throttled in themanner of a liquid engine. Reverse hybrids, using a solidoxidizer (e.g., ammonium perchlorate) and a liquid fuel(hydrazine) have been tested. Since solid oxidizers are notas energetic as liquid oxidizers, there is little value in thereverse hybrid concept.

The Firebolt (AQM-81), a target vehicle, uses IRFNAas the oxidizer with a PBAN based solid fuel. A typicalfuel composition is

1. 80 wt% polybutadiene/acrylic acid/acrylonitrile2. 20 wt% polymethylmethacrylate (Plexiglas

❤™)

A 20,000 lb f (88.96 kN) hybrid unit utilizing 90% hy-drogen peroxide and an aluminum-enriched polyethylenegrain was successfully tested. The oxidizer was catalyt-ically decomposed, producing hot gases that ignited thefuel in a smooth and reliable manner.

The solid lithium/ClF3 concept was discussed in Sec-tion III.D.5. Experimental grains consisting of powderedlithium held in a solid binder have been made and tested.Because lithium is very light, it has been difficult to gethigh metal loadings in a usable grain.

American Rocket Company (AMROC) attempted alaunch from Vandenberg AFB with a hybrid rocket ve-hicle. The oxidizer was LOX and the fuel was a butadienewith the consistency of a soft automobile tire.

I. High Energy Density Materials

High-Energy Density Material (HEDM) research has beenunder way since the mid-1980s. Rather than presumewhich of the previous categories of propellants will beembellished with the development of these new technolo-gies, the area will be discussed separately. Fundamentally,the potential for these special molecules is that they con-tain and will release energy far beyond what is possiblewith normal combustion. The strained ring hydrocarbonsdiscussed in Section III.D.1 as kerosine additives are ac-tually part of the HEDM activity with near-term potential.

The following are research areas with potential payofffor improved performance rocket propellants:

1. Polymeric or poly nitrogen (N4, N6, N8). In 1998,AFRL announced that scientists at Edwards AFB hadidentified N5 as the cation combined with the AsF6

anion. The cation’s heat of formation was calculatedat 353 kcal/mol

2. Polymeric or poly oxygen (O4, O6, O8)3. Hydrides of boron (BH2, BH4)4. B-N analogs of prismane (B3N3H6)5. Atomic additives to hydrogen in an O2/H2 system.

Adding energetic atoms to a LOX/LH2 system has thepotential to increase specific impulse by the followingamounts:

1. Carbon atoms + 49 sec2. Boron atoms + 32 sec3. Aluminum atoms + 27 sec4. Hydrogen atoms + 19 sec

The use of F2/H2 in a vehicle also has a 19-sec increasein specific impulse over LOX/LH2, but it is hoped thatthe new technology will be feasible without the safetyproblems and complications of using liquid fluorine.

IV. HANDLING AND STORAGE OF LIQUIDROCKET PROPELLANTS

This section provides additional information on liquidpropellants in current or recent use or that have beenthe subject of considerable evaluation and could be se-lected for future propulsion systems. Rocket propellantsare by nature hazardous materials and must be handled

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and used only in appropriate facilities by personnel whohave the proper training and knowledge of their propertiesand approved personal protective equipment. The follow-ing outlines the type information available and the generalprecautions needed, but is insufficient to provide the depthof instruction required by personnel. Current regulations,operating procedures, and technical information must beconsulted for the safe use of propellants.

A. Ammonia

Liquid anhydrous ammonia is a commercial chemical,used for refrigeration and as a fertilizer. It is very stableand is not shock sensitive. Being alkaline and a reducingagent, ammonia will react with most acids. It is soluble inwater, alcohol, and many other solvents.

1. Specifications

Propellant-grade ammonia procured under SpecificationO-A-445 contains 99.8 wt% NH3. The major impurity iswater that is limited to a maximum of 0.5 wt% to reducecorrosion. Excessive oil will deposit in engine passagesand foul igniters, thus it is controlled to 5.0 ppmw (partsper million weight) maximum. The purity of ammoniaproduced in the United States is consistently high. In theamounts normally present in refrigeration grade ammonia,pyridine, naphthalene, and hydrogen sulfide were foundnot to adversely affect rocket engine performance.

2. Material Compatibility

Anhydrous ammonia will not corrode iron, steel, or alu-minum. Since it is difficult to avoid some moisture in op-erational systems, copper, zinc, and their alloys should notbe used. The use of nickel and stainless steel (300 and 400series) is recommended.

Suitable nonmetallic materials for anhydrous ammoniaservice are tetrafluoroethylene polymer (Teflon

❤™), chlo-rotrifluoroethylene (Kel-F

❤™), ammonia-resistant rubber,and asbestos (free of graphite and grease).

3. Safety Considerations

Anhydrous ammonia must be handled with caution be-cause it is a toxic, flammable compound shipped andstored under pressure. The principal hazards to person-nel are inhalation of the vapor and contact of the liquidand vapor with skin and eyes. Liquid ammonia producessevere burns on contact due to its caustic action and lowtemperature. Gaseous ammonia is a strong irritant andcan injure the eyes and respiratory tract. The TLV is 50ppm. The flammability range of ammonia in air is 16 to

27 vol%. Its flash point is <32◦F (273 K) and autoignitiontemperature is 1204◦F (924.3 K).

Water fog effectively suppresses ammonia fires since itwill cool the burning surface and reduce ammonia vaporpressure by absorption and dilution.

4. Storage and Handling

Anhydrous ammonia is transferred and stored as a lique-fied, compressed gas and is relatively easy to handle withreadily available equipment designed for ammonia ser-vice. Storage and transfer areas must be well ventilated,free from oxidizers, flammable materials, excessive heat,and sparks.

Ammonia may be stored in shipping cylinders or in non-refrigerated bulk storage tanks. Care must be taken to pre-vent overfilling containers. A tank is considered overfullwhen at any time there is insufficient vapor space abovethe liquid to provide adequately for the thermal expansionof the liquid as ambient temperatures fluctuate.

5. Transportation

Commercially available cylinders, rail tank cars, and cargoor portable tanks of various sizes that meet U. S. Depart-ment of Transportation (DOT) specifications are used forshipping anhydrous ammonia. DoT classifies ammonia asa nonflammable gas for shipments because the concen-tration range of flammability is narrow and the ignitiontemperature is high.

B. Chlorine Pentafluoride

Chlorine pentafluoride (ClF5) is a colorless gas under nor-mal conditions. It can be maintained as a liquid underits own vapor pressure, 58.8 psia (405.2 kN/m2) at 77◦F(298.2 K). CLF5 is a very strong oxidizing agent that ishypergolic with common rocket fuels and will react vigor-ously with water and most combustibles. It is not sensitiveto heat, shock, or sparks.

The information presented in this section also appliesto chlorine trifluoride (ClF3) and similar interhalogen ox-idizers.

1. Specifications

The principle impurity in ClF5 and ClF3 is hydrogen fluo-ride (HF), which causes corrosion and particulate forma-tion. The typical assay is product: 99.8 wt% minimum andHF: 0.2 wt% maximum. The procurement specifications,MIL-P-27413 (ClF5) and MIL-P-27411 (ClF3), were re-cently canceled.

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In a clean, dry, passivated system, ClF5 is compatible witha wide range of metals. However, at elevated temperatures,it will react with them. It is important not to have anysources of ignition, such as contaminants, metal-to-metalfriction (as in a valve), or mechanical impact sufficientto deform metal. The following metals have been success-fully used: 1160, 3003, 6061, 2024, 5052, 6063 aluminum;Inconel

❤™, Monel❤™, nickel; stainles steel, all; and copper,

brass, and carbon steel.The following metals should not be used in contact with

ClF3 or ClF5: titanium, tantalium, niobium, and tungsten.There are no elastomers or plastic materials suitable for

liquid flow or high pressure systems. For low gaseous flowrates where pressures are under 400 psig (2.76 MN/m2)or static liquid systems, polytetrafluoroethylene (e.g.,Teflon

❤™, Halon, TFE) may be used. These materialsmust be properly cleaned and dried, and no hydrocarbon,alcohol, or halogenated solvents used. The solvents canpenetrate into the plastic and are impossible to removecompletely.

The use of all lubricants, including fluorocarbons, areprohibited.


While they do not burn or explode by themselves, ClF3 andClF5 must be considered fire hazards due to their ability toignite most materials. The TLV-C for ClF3 is 0.1 ppm: noofficial value has been established for ClF5, but 3.0 ppmhas been suggested.

4. Cleaning and Passivation

All equipment in contact with the interhalogens must bescrupulously cleaned and passivated. All grease solvents,scale, weld spatter, and moisture must be removed. Allcomponents must be completely disassembled to assurethat no dead areas containing the slightest residue areoverlooked. Every square inch of the system must be pas-sivated with a tenacious fluoride coating that protects thesurface against further attack by the interhalogen. Passi-vation is accomplished by exposing the system to a mix-ture of fluorine and nitrogen or helium gas. The partialpressure of fluorine is gradually increased until the sys-tem is 100% filled with gaseous fluorine at the work-ing pressure of the system. The system may get veryhot and even burn through if the pressure of fluorine israised too quickly. Even after the system was passivated,these propellants have been known to burn through as theyflow through any tight turns such as a standard elbow.Only long radius tube bends will turn the flow of thesepropellants.

5. Handling and Storage

ClF5 is stored and transferred as a liquefied gas. Tanksshould meet the ASME code for unfired pressure ves-sels. Remote-operated, packless-type valves with metal-to-metal seats should be used. Suitable diaphragm-typepumps are commercially available for the interhalogens.Soft copper and 1100 aluminum may be used for gasketsat flanged fittings; however, all-welded connections arerecommended.

6. Transportation

The interhalogens are shipped in mild steel cylinders meet-ing DOT specifications. The largest containers presentlyapproved have a product capacity of 2000 lb (907 kg).

C. Fluorine

Fluorine is the strongest oxidizing agent and one of themost reactive substances known. Under certain conditions,it can react with most elements and compounds, the excep-tions being some of the rare gases and completely fluori-nated compounds. If kept free of moisture, which reacts toform HF, fluorine is noncorrosive to clean and passivatedmetals. Liquefied fluorine is hypergolic with all rocket fu-els. Gaseous fluorine has a yellowish color, and the liquidis amber. It has a very pungent odor. Fluorine is stable andliquid is completely miscible with liquefied oxygen andnitrogen.

1. Specifications

Fluorine procured under MIL-P-27405 has a minimum as-say of 99.0 vol%. Oxygen and nitrogen are the major im-purities (0.8% maximum) with some CF4 and CO2 (0.2%maximum) present. HF is the most troublesome impurity,causing container corrosion and forming particulate mat-ter. HF is limited to 0.5% maximum, with 0.2% a typicalvalue. In liquefied fluorine, HF is a solid and can be re-moved by filtration. Ozone and peroxides, formed duringthe electrolytic manufacturing process, are removed at theplant.


Materials selected for fluorine service must be resistantto attack by fluorine and possess appropriate strength atcryogenic temperatures.

The following metals are recommended for use in liq-uefied fluorine systems: nickel, aluminum, copper, brass,Monel

❤™, 304L, and 316 and 321 stainless steel. In addi-tion to these, mild steel is satisfactory for gaseous fluorinefrom 20 to 160◦F (244 to 344 K).

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There are no plastic or elastomeric materials accept-able for liquefied fluorine under flow conditions or forhigh-pressure gaseous fluorine. Under moderate flow ratesand low pressure polytetrafluoroethylene (Teflon

❤™) maybe used. It is very important to assure that clean, highquality Teflon

❤™ (no metallic impurities from the extru-sion process) is used. Since it is very difficult to removesolvents that may have impregnated the plastic, solventcleaning should not the used.


Fluorine is very toxic (TLV 1.0 ppm). Contact with thevapor will cause some irritation and burns. Liquefied flu-orine will destroy body tissues on contact. Under spillconditions, liquefied fluorine will react with essentiallyall materials, usually causing a fire; thus it must be calleda fire hazard. Uncontrolled mixing of liquefied fluorineand rocket fuels as in a spill situation will produce severedeflagrations.

4. Cleaning and Passivation

Proper cleaning and passivation of fluorine systems is thekey to successful operation. During the passivation pro-cess, a tightly adhering film of metal fluoride is formedon the interior surfaces to protect the metal from furtherfluorine attack. Components should be disassembled andplastic parts removed. All contaminants, especially hydro-carbons, must be removed with solvents and detergents.Scale, slag, and weld beads are removed by sand blastingand acid pickling. Deionized or distilled water is used as afinal rinse, and the equipment dried with heated nitrogengas. The system is reassembled and purged with nitrogen.

Passivation is accomplished by slowly replacing thenitrogen with gaseous fluorine at atmospheric pressure.After this nitrogen is replaced, the pressure is increasedslowly to the operational pressure and left under observa-tion for about two hours. If there is no internal heating,the system is considered passivated.


Liquefied fluorine is always kept in a closed system undera low positive pressure to avoid release of its vapors tothe atmosphere. Refrigeration is required that is usuallysupplied by the boil-off of liquefied nitrogen. Specialtanks are used and constructed with three horizontal,concentric shells. The inner shell holds the product,an intermediate shell contains liquefied nitrogen, andthe outer shell serves as a vacuum jacket. Monel

❤™ isthe preferred material for the two inner shells, althoughstainless steel may be used if proper consideration is given

to low temperature strength. The outer vacuum jacketmay contain insulation such as perlite or Santocel

❤™. Aslong as the liquefied nitrogen is replenished, the fluorinewill keep indefinitely in the liquid state.

No pressure relief devices are to be inserted in theinner fluorine tank. The liquefied nitrogen jacket willbe equipped with vents and a pressure relief systemtypical of liquefied nitrogen systems. Remote-operated,packless-type valves designed for fluorine service shouldbe used. All-welded connections are recommended andthreaded connections should be avoided. Flanges or com-pression fittings require soft copper or aluminum gaskets,rings, or seals. Liquefied fluorine has been pumped satis-factorily; however, suitable pumps are not generally avail-able. Transfer can be accomplished by using 1–5 psi (7–35 kN/m2) differential helium pressure.

6. Transportation

Liquefied fluorine is shipped in tank trucks that hold5000 lbs (2268) kg/lbs of product. Special DOT permitsare required These truck trailers are of the three concen-tric shell design described in Section IV.C.4. The liquefiednitrogen volume is sufficient to last at least 10 days with-out replenishment, more than adequate for cross-countryshipments.

D. Hydrazines

The widely used hydrazines, anhydrous hydrazine, MMH,UDMH, and the 50% hydrazine/50% UDMH blend(Aerozine-50) are clear, colorless, oily liquids with dis-tinctively fishy ammoniacal odor. These strong reducingagents are completely miscible with water, alcohols, andmost solvents. They are very hygroscopic, and tend toabsorb carbon dioxide and oxygen from the atmosphere.These fuels are hypergolic with the common storablerocket oxidizers. Chemically, hydrazine is the most re-active of the family, and extra care must be exercised inhandling. All are stable to friction and shock but are ther-modynamically unstable. Iron, nickel, cobalt, copper, andplatinum metals and their oxides will catalytically decom-pose hydrazine and MMH.

1. Specifications

The military specifications covering these fuels arehydrazine, MIL-P-26536; MMH, MIL-P-27404; 50–50blend, MH-P-27402; and UDMH, MIL-P-25604. Typi-cally the fuel assay is 98 wt% or better, with the majorimpurities being water and other amines. Impurities in hy-drazine can adversely affect its stability and can damagecatalysts when it is used as a monopropellant. Thus for the

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monopropellant grade hydrazine the following contami-nants are controlled: carbon dioxide, 30 ppm maximum;aniline, 0.5 wt% maximum; iron, 20 ppm maximum; chlo-ride, 5 ppm maximum; nonvolatile residue, 50 ppm max-imum; and others (MMH, alcohol), 200 ppm maximum.


The hydrazines are noncorrosive to most metals. Becauseof catalytic effects, only known materials should be used.Recommended metals and alloys include: 304, 304L, 316,347, and 17–7PH stainless steel; 1100, 2014, 2024, 2219,5052, and 6061 aluminum; 6A1–4V titanium. The follow-ing metals should be avoided: cadmium, cobalt, copperand alloys, iron, lead, magnesium, mild steel, and zinc.A limited number of elastomers and plastics are suit-able for hydrazine family use: AF-E-332; butyl rubber;ethylene-propylene copolymer; and polytetrafluoroethy-lene. Viton

❤™ and chlorotrifluoroethylene should not beused except for short exposures.

No lubricant is completely satisfactory. The followinghave been used: Apiezon L, Dow Corning Silicon Com-pound 11, Fluorolube GR-470, Kel-F grease, Krytox PR-240AC, and Ready Lube 200.


The hydrazines are toxic and and will irritate eyes andskin. They can be absorbed by inhalation, through theskin, and by ingestion. The TLVs are hydrazine; 0.1 ppm;MMH, 0.2 ppm; and UDMH, 0.5 ppm. They are sus-pected carcinogens. The hydrazines will not explode, butthe vapors will make explosive mixtures with air. Theyare flammable in a broad range of concentrations in air:hydrazine. 4.7–100% MMH, 2.5–98%; and UDMH, 2.0–90%. Spills of these fuels on rags, dirt, rust, sawdust, andso on can result in spontaneous ignition. Because of theease of ignition, it is common in the vicinity of an opencontainer or a spill to have tongues of flame appear. Thisphenomenon is more apparent at night.


Storage tanks with capacities up to 20,000 gal (75.7 m3)and meeting the ASME Code for Unfired Pressure Ves-sels are used for the hydrazines. Type 304 stainless steelis commonly used. For small volumes, the 55-gal (0.2 m3)shipping drums may be used for storage. All containersmust be thoroughly cleaned and inspected to assure thatno catalytic materials are present. All air is purged fromcontainers and piping with an inert gas, usually nitrogen,prior to filling with fuel. An inert gas blanket of a slightpositive pressure is maintained at all times in systems to

keep out air and moisture. Transfer of fuels may be ac-complished with pumps or differential pressure, and ar-eas must be free of organic and oxidizing materials andrust. There must be no open fires or electrical sparkspresent although as noted above, spontaneous ignition iscommon.

Hydrazine has been observed to undergo explosive de-composition by rapid adiabatic compression of vaporswhen subjected to high hydrodynamic pressure (waterhammer).

5. Transportation

The hydrazines are shipped in DOT-approved tank trucks,rail cars, portable tanks, and 55-gal (m3) drums. Red“Flammable Liquid” and white “Poison” labels are re-quired.

E. Hydrogen

Liquefied hydrogen is transparent, colorless, and odorless.Because of its very low boiling point, release of liquefiedhydrogen in a normal atmosphere will result in a visible,voluminous cloud of condensed moisture. Hydrogen isnoncorrosive, but some metals are subject to hydrogen em-brittlement, especially at higher temperatures. Hydrogengases are combustible with air over the range of 4–75 vol%at 68◦F (293 K). Mixtures of hydrogen with air from 18to 59% by volume will detonate. Liquefied or gaseoushydrogen is hypergolic with liquefied and gaseous fluo-rine, chlorine trifluoride, and chlorine pentafluoride, andwill form combustible mixtures with other oxidizers. Allsubstances are essentially insoluble in liquefied hydrogen.Helium is slightly soluble, up to ∼1% under pressure.

Liquefied hydrogen is chemically stable, but it is physi-cally stable only when stored under suitable conditions. Ina properly designed container, the evaporation rate can beas low as 1.5% per day in a 1000-gal (3.79 m3) container,and less in larger vessels.

1. Specifications

Liquefied hydrogen conforming to MIL-P-27201 has anassay of 99.995 vol% H2 minimum and a minimum con-version to 95.0 vol% parahydrogen. Oxygen and hydro-carbon gases are limited to a maximum of 1.0 ppmv (partsper million volume) each. Up to 48.0 ppmv of inert gases(e.g., He, N2) is permitted.


A prime consideration for materials selection for lique-fied hydrogen service is the ability to retain satisfactory

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physical properties at the liquefied hydrogen boiling point.Materials must be able to withstand the high thermalstresses imposed by large and rapid temperature changes.Many iron alloys, for example, lose their ductility and be-come brittle.

Recommended metals for liquefied hydrogen serviceinclude: 300 and other austenitic series of stainlesssteel; copper, brass, and bronze; aluminum; Monel

❤™; andEverdur

❤™.All plastics and elastomeric materials lose their flexi-

bility at liquefied hydrogen temperatures. Some materi-als successfully used in applications are polyester fiber(Dacron

❤™); tetrafluoroethylene (Teflon❤™ or equivalent);

chlorotrifluoroethylene, unplasticized (Kel-F❤™); nylon;

Mylar❤™; and asbestos impregnated with TFE.

Lubricants solidify at liquefied hydrogen temperaturesand are not generally used. In some special applications,vacuum grease can be used as a sealant with O-rings.


Hydrogen is extremely flammable. Unless carbonaceousimpurities or other contaminants are present, the flamemay be invisible. Explosive hazards exist when frozen air(oxygen) collects in liquefied hydrogen, and when gaseoushydrogen is mixed with air in a confined space. Skin con-tact with liquefied hydrogen or with equipment containingit can cause severe frostbite. Air in contact with lines andtanks can liquefy and produce an oxygen-rich environ-ment. Hydrogen is an asphyxiant, but it is not toxic.


Liquefied hydrogen may be stored in stationary or mobiletanks of appropriate design and materials and tested in ac-cordance with ASME or DOT specifications for pressurevessels. The tanks are vacuum insulated, and evacuatedperlite, Santocel

❤™, or multilayer super insulation is usuallyused. Liquefied hydrogen systems are kept under positivepressure to prevent entry of air. The internal structure sup-porting the inner hydrogen tank within the vacuum shellis made as light as possible to reduce thermal leaks. Be-cause the density of liquefied hydrogen is very low, thisinternal structure is considerably lighter than for tanksdesigned for other commodities. This weight restrictionmust be observed when liquefied nitrogen is used for pre-cooling and purging or if the container is used for anotherproduct.

Liquefied hydrogen systems must be thoroughlycleaned to remove all oils and surface films, rust, weldscale, and other particulate matter and then dried to elimi-nate all moisture. Prior to use, the system must be purgedwith warm hydrogen or helium gas to remove all traces

of air. Only vaporized gas from liquefied hydrogen shouldbe used; the impurities present in most cylinder H2 makeit unsuitable for this purpose. All lines should be vacuumjacketed or insulated.

5. Transportation

DOT classifies liquefied hydrogen as a flammable com-pressed gas. Rail tank cars, tank trucks, and small cylin-ders holding about 200 lb (91 kg) of product are used forshipping liquefied hydrogen. These containers are vacuumjacketed and insulated and are equipped with pressure re-lief devices. A red “Flammable Gas” label is required.

F. Hydrogen Peroxide

Concentrated hydrogen peroxide is both a strong oxidiz-ing agent and a monopropellant. It is a clear, colorlessliquid that is slightly more viscous than water. Some gasevolution can be seen. It can be rapidly decomposed uponcontact with many inorganic compounds, including ironoxide, silver, and potassium permanganate. This contactyields water (steam), oxygen, and heat. Hydrogen perox-ide will not burn, but will support combustion of many or-ganic materials. The exothermic decomposition reactionreleases sufficient heat to ignite combustible materials. Itis hypergolic with amine fuels, including the hydrazines,and the metal hydrides. It is considered hypergolic withhydrocarbon fuels; however, the heat release rate is in-sufficient to provide reliable ignition in a rocket enginewith out the use of some other ignition system. For thisreason, a two-stage engine concept is used in which theperoxide is first catalytically decomposed and the hotgases injected into the rocket engine chamber with thefuel.

1. Specifications

Hydrogen peroxide conforming to MIL-P-16005 (recentlycancelled) has an assay of 90.0–91.0 wt%. Impurities thatpromote decomposition must be avoided if good stor-age life is to be obtained; however, stabilizers or catalystpoisons cannot be tolerated if catalyst packs are not tofail prematurely. For this reason, specific impurities areclosely controlled to provide reasonable storage stabil-ity and to avoid poisoning of catalysts. Dissolved metalscan cause homogeneous decomposition. Hydrogen perox-ide stability is determined by holding a sample at 100◦C(373.15 K) for 24 hr. A maximum of 5% active oxy-gen loss is permitted. Aluminum content is controlledto 0.5 mg/1 (g/m3). Chloride was found to cause corro-sion of containers, thus it is controlled to a maximum of1.0 mg/1 (g/m3), and 3–5 mg/1 (g/m3) of nitrate is added as

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a corrosion inhibitor. Phosphate is a strong catalyst poisonthat enters during manufacture. Since it cannot be entirelyeliminated at a reasonable cost, 1.0–4.0 mg/1 (g/m3) tin(as stannous chloride) is added to offset its effects. Sul-fate, also a catalyst poison, is limited to 3.0 mg/1 (g/m3)maximum. Extensive studies have shown that up to 200ppm carbonaceous materials (residuals from the principalmanufacturing processes) have no detrimental effect onstability, catalysts, or engine performance.


In selecting materials for hydrogen peroxide equipmentand systems, the effect that the hydrogen peroxide willhave on the material is much less important than the effectthe material may have on the peroxide. Only materials thathave been found compatible by actual test should be used.

Aluminum 1060, 1260, 5254, and 5652 are suitable forlong-term contact with hydrogen peroxide, as in storagetanks and related piping.

In general, 300-series stainless steel (347 preferred) andaluminum 1100, 3003, and 6061 are usable for rocket en-gine components and test facilities where the contact withhydrogen peroxide does not exceed four weeks a normaltemperatures or a few hours at about 160◦F (344 K).

Some other common aluminum alloys and stainlesssteels may be used when the contact is less than one hourprior to use, as in engine valves, feed lines, and injectors.

Of the nonmetals, only polytetrafluoroethylene(Teflon

❤™) and chlorotrifluoroethylene (Kel-F❤™) are the

only plastics suitable for extended contact with hy-drogen peroxide. A number of materials, including poly-ethylene, selected silicone rubbers, Viton

❤™ A, and poly-vinyl chloride, can be used for limited exposures. Cautionmust be exercised since the quality of these products varygreatly and not all samples are safe to use. The amountand type of plasticizers and previous exposure to solventscan have a profound effect of compatibility.

Chlorofluoro oils and greases (e.g., Halocarbon❤™,

Kel-F❤™) are suitable for limited use with hydrogen per-

oxide but should not be used with aluminum.


Hydrogen peroxide may ignite combustible materials oncontact, including dry vegetation and ordinary clothing.Contamination, such as rust, dirt, metal filings, and soon, can cause rapid decomposition, releasing oxygen andconsiderable heat. The evolution of oxygen and steam mayoverwhelm any venting system causing the container torupture.

Hydrogen peroxide will irritate skin, eyes, and the res-piratory system. Extended contact will cause burns. While

it has a TLV of 1 ppm, toxicity is generally not a problemdue to its low vapor pressure.

4. Cleaning Procedures

All containers, components, and surfaces that are sub-ject to exposure to hydrogen peroxide must be clean andpassivated. All equipment must be disassembled and alu-minum, stainless steel, and plastic parts separated. Metalcomponents are degreased with trichloroethylene (or otherapproved solvents or detergents) and then rinsed with al-cohol and then with distilled or deionized water. Stain-less parts are pickled with 70% nitric acid, and aluminumparts are treated with a 25% sodium hydroxide solutionfollowed by a 45% solution of nitric acid. After a thoroughrinse with deionized water, the components are passivatedby filling large equipment and immersing small parts in70–90% hydrogen peroxide for 4 hr and the decompositionrate of the hydrogen peroxide noted. If extensive decom-position is observed, the part is recleaned or discarded.Plastic materials are cleaned and passivated in a similarmanner; however, it is best to degrease with detergentsrather than solvents.


Horizontal-type aluminum storage tanks are normallyused for hydrogen peroxide. All storage tanks must bevented and a vent filter used to prevent dirt from enteringthe tank. The 30-gal (0.11 m3) aluminum shipping drumsare suitable for storage containers when small quantitiesare involved. Shielding of the containers from the rays ofthe sun is desirable. Good housekeeping is essential tominimize fire hazards and to provide a safe work area.All storage and transfer areas, as well as surrounding ar-eas that may be reached by leakage or spills, should becleared of organic matter and combustibles. All contain-ers, piping, and equipment that will be in contact withhydrogen peroxide must be of appropriate materials andthoroughly cleaned and passivated before use. A wide va-riety of valves, pumps, gauges, and transfer equipment arecommercially available, but care must be exercised in theirselection. Transfer of hydrogen peroxide may be accom-plished by pumping, by vacuuming, or by pressuring withclean, dry nitrogen.

6. Transportation

DOT classifies hydrogen peroxide as a corrosive liquidand requires a white label during shipment. Rail tank cars,up to 8000 gal (30.28 m3); tank motor vehicles, 4000 gal(15.14 m3); and 30-gal (0.11 m3) drums that meet DOTspecifications are used for hydrogen peroxide shipments.

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Two types of drums are available: a double-head drumsuitable for all freight shipments and a single-headdrum approved only for full carload and full truckloadshipments.

Hydrogen peroxide is commonly shipped at 70% con-centration as a feed stock for manufacture of semicon-ductors. Uses as a rocket oxidizer are above 90%. X-LSpace Systems offers on-site concentrators to avoid ship-ment of the higher concentrations. Degussa-Hus Inc. willship concentrations up to 87.5%.

G. Kerosine

Propellant-grade kerosine, formerly called RP-1, is a high-boiling hydrocarbon fraction. It is a clear liquid, with aslight red color from an added dye. It is very similar to JP-4turbine engine fuel but with a controlled aromatic content.It is chemically stable and is insensitive to shock. Kero-sine, like gasoline and diesel and jet fuel, reacts only withstrong oxidizers or under high temperatures and pressures.High temperatures, especially in the presence of air canaccelerate the formation of gum and sediment. Like otherhydrocarbons, kerosine is flammable, and its vapor willform explosive mixtures with air. It has a fairly low vaporpressure. It is insoluble in water, and it is an excellientsolvent and is soluble in many organic solvents.

1. Specifications

The major characteristics of kerosine defined by MIL-P-25576 are distillation, 10% point, 365–410◦F (458.2–483.2 K); end point, 535◦F (552.6 K) maximum; residue,1.5 vol% maximum; gravity, 42.0–45.0◦ API; aromatics:5.0 vol% maximum; olefins, 1.0 vol% maximum; and par-ticulate: 1.5 mg/1 (g/m3) maximum.


A wide range of ferrous and nonferrous alloys is suitablefor storage and shipping containers and associated com-ponents.

Only petroleum-resistant plastics and elastomers are tobe used with kerosine. Some suitable nonmetals are cork,fluorocarbons, asbestos, polyvinylchloride, nitrile rubber,Buna

❤™-N, Neoprene ❤™, and polyethylene.

Graphite-based, molybdenum disulfide, some silicon,and fluorocarbon lubricants may be used. Kerosine willdissolve petroleum-type lubricants.


Kerosine vapors will form ignitable mixtures with air, asdoes gasoline and diesel fuel. While it will ignite upon

contact with fluorine and chlorine trifluoride, it is not hy-pergolic with most rocket oxidizers. Kerosine will reactwith nitric acid, hydrogen peroxide, and strong oxidizerswith evolution of heat and may ignite. Such mixtures ofkerosine and oxidizers are likely to be sensitive to shock,heat, and sparks. The flash point is 110◦F (316.5 K). Kero-sine is a moderate skin irritant; scaling and fissuring ofskin can result from frequent contact. Inhalation of vaporscan cause narcosis. The toxicity of kerosine (TLV) is afunction of the aromatic content, e.g., benzene.


Huge quantities o f hydrocarbon fuels are used through-out the world. The same equipment and procedures maybe used for propellant-grade kerosine, with due precau-tions taken to minimize corrosion and contamination. Fireprotection is the major concern in facility design. All sta-tionary and mobile tanks should be grounded, electricaldevices must be explosion proof, and all vents protectedwith flame arresters. Buildings are to be constructed offire-resistant materials, and the area kept free of all vege-tation and combustible debris.

5. Transportation

Kerosine is classified by DOT as a “Flammable Liquid.”Standard petroleum rail cars and tank trucks are commonlyused for shipping.

H. Nitric Acid

Although fuming nitric acid is a commercial chemical,special formulations are used as rocket oxidizers. Whitefuming nitric acid (100% HNO3) is unstable, dissociatingto NO2 and water [Eq. (14)], and very corrosive, attackingmost metals. The addition of NO2, HF, and water greatlyimproves nitric acid storability, making it a valuable pro-pellant. White fuming nitric acid is colorless, but becomesdeep yellow to reddish brown as NO2 forms. The currentrocket propellant grades are a dense reddish brown. Fum-ing nitric acid is hypergolic with the hydrazines and mostamine fuels. It will react with water and many organicmaterials with evolution of heat and NO2 gas.

1. Specifications

Current propellant-grade nitric acid formulations, pre-sented in Table X, are procured under specification MIL-P-7254. The composition of Type IIIA was selected to pro-vide the greatest stability and minimum corrosion. TypeIIIB has a lower aluminum salt content otherwise it isidentical to Type IIIA. The lower water content of Type

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IIILS (for limited storability) trades storability for a littlehigher performance, again for a special application wherelong storage life is not important. Type IV has the highestspecific impulse and density, but is more corrosive.


Aluminum is the material of choice for fuming nitric acidwhen maximum storability and minimum metal attackis needed. The following aluminum alloys are recom-mended: 1060, 1100, 3003, 3004, 6061, 5052, and 5154.

Fabrication of tanks and equipment from stainless steelrequires care. Crevices that are hard to clean should beavoided. To avoid loss of corrosion resistance, weld areasin unstabilized grades must be annealed to redissolve thechromium carbides and rapidly quenched. These stainlesssteel alloys have been found satisfactory: 303, 304ELC,316, 347, 321, 19–9DL, and Durimet

❤™ 20.Fuming nitric acid severely corrodes copper and its al-

loys and reacts violently with titanium.Polytetrafluoroethylene and polychlorotrifluoroethy-

lene are the only nonmetals resistant to fuming nitric acid.The perfluorocarbon type lubricants are also satisfactory.


Fuming nitric acid will not burn and is stable to impact,sparks, and shock, but it may form explosive mixtures withkerosine, other hydrocarbons, and alcohols. It will reactwith many organic materials and may ignite them. In animproperly prepared container or at elevated temperatures,internal pressure buildup may occur. The TLV for nitricacid is 2 ppm. Nitric acid fires and spills can be controlledwith copious amounts of water. The residues then shouldbe neutralized with sodium carbonate. Water will cause anexothermic reaction with nitric acid, causing the evolutionof large amounts of NO2; however, high-pressure waterfog will contain the fumes. A special water-based foamconsisting of acrylic surfactants and pectin is also availablefor fume suppression.


Aluminum and stainless steel tanks meeting ASTM Boilerand Pressure Vessel Codes, equipped with pressure reliefdevices, are used for fuming nitric acid storage, with alu-minum 11,000 gal (41.64 m3) tanks being the most de-sirable. The 55-gal (0.2 m3) shipping drums may also beused for storage if in a covered area. Storage areas mustbe free of all vegetation and combustible material and besurrounded by a dike. Tanks and components in contactwith acid must be thoroughly cleaned, degreased, deter-gent cleaned, and all scale and weld slag removed. Stain-

less steel components are also pickled with a dilute nitricacid/HF solution. Precautions must be taken to avoid gal-vanic corrosion when dissimilar metals are used in a sys-tem. Chemical type pumps are used for nitric acid transfer.Pressure transfers are not recommended.

5. Transportation

Only aluminum containers are used for cargo truck, railcar, and drum shipments of fuming nitric acid. An “Oxi-dizer” and a “Poison” label are required by DOT.

I. Nitrogen Tetroxide

Nitrogen tetroxide is the most widely used storable liquidoxidizer. N2O4 is in temperature equilibrium with NO2. Itis the NO2 that produces the dense reddish-brown colorassociated with this oxidizer in both the liquid and gaseousforms. As the temperature is lowered, the equilibriumshifts toward N2O4 and the color changes from reddishto straw tan to a very pale yellow. At the normal boilingpoint (70.1◦F; 294.3 K), the mixture contains about 15%NO2 and essentially 100% at 302◦F (423.2 K). Near thefreezing point, the liquid becomes colorless. The MONs,mixed oxides of nitrogen, contain NO with the numericaldesignation representing the nominal NO content. Com-positions MON-1 and MON-3 contain NO as a stress cor-rosion inhibitor, and composition MON-10 and higher as afreezing point depressant. The nitric oxide complexes withthe NO2 to form N2O3, a blue substance, which imparts agreenish color to the nitrogen tetroxide; hence the terms“green” N2O4 (for the MONs) and “red-brown” N2O4 (nonitric oxide) are frequently used. MON-1 is the commonlyused composition. MON-3 is specified for the Space Shut-tle to assure that throughout handling, servicing, and flightoperations, the NO content remains above the 0.4% neededfor stress corrosion protection. Neat N2O4 (NTO) was usedin the Titan II missile.

MONs with the higher levels of NO will have highervapor pressure. Otherwise, the safety, material compati-bility, and handling characteristics for all of the MONs aresimilar and are treated alike in this section.

1. Specifications

MIL-P-26539 covers the procurement of N2O4 (gradeNTO), MON-1, and MON-3. The requirements are pre-sented in Table XVI. MON-15 and MON-25 are also pro-cured under this specification.


Moisture-free N2O4 and MONs are noncorrosive. The fol-lowing metals are suitable for most purposes: aluminum,

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TABLE XVI Nitrogen Tetroxide Specification Requirements(Chemical Composition Limits)a

Requirement NTO MON-1 MON-3

N2O4 assay (wt%) 99.5–100.0 — —

Nitric content (wt%) — 0.6–1.0 2.5–3.0

Water equivalent (wt%) 0.17 0.17 0.17

Chloride (wt% max.) 0.040 0.040 0.040

Particulate (mg/1 max) 10 10 10

Iron (ppm max.) — — 0.5

a From National Aerospace Standard 3620.

carbon steel, Inconel ❤™, nickel, stainless steels, and tita-

nium. N2O4 and the MONs react readily with moisture toform acids. For this reason the use of titanium or 300-seriesCRES are highly recommended for long-term storage andpropellant-feed systems. Of the non-metals, the followinghave been found satisfactory: asbestos, cotton and oil-free;AFE-124R elastomer; ceramic, acid-resistant; polyte-trafluoroethylene: and polychlorotrifluoroethylene. Flu-orocarbon lubricants and Teflon

❤™ tape may also be used.


N2O4(NO2) is a toxic gas, with a TLV of 3 ppm. It re-acts exothermically with water to form nitric and nitrousacids, with an evolution of NO2 gas. The MONs are hyper-golic with the hydrazines and most amine fuels. They willreact with many organic materials and may form explo-sive mixtures with hydrocarbon fuels and partially halo-genated solvents. While these oxidizers will not burn, theywill support combustion. The MONs are thermally stableand are not sensitive to mechanical impact. They will notexplode.

Because of the volatility of the MONs, most leaks andspills will result in the formation of vapors. Water addedto a pool of N2O4 will greatly increase vapor evolution;however, high-pressure water fog will suppress the fum-ing, but must be applied until the N2O4 liquid is suffi-ciently diluted. Special foams have been developed thatare effective in suppressing the vapors for about 10 min.


The nitrogen oxides may be stored in mild steel tanksconforming to the ASME Boiler and Pressure VesselCode. Shipping cylinders are also used for storage con-tainers. Pressure relief devices are required. Because smallamounts of water will form nitric acid and cause corro-sion, the trend is to use stainless steel tanks and transferequipment. Iron corrosion products are very insoluble in

the MONs and will rapidly clog filters. The small amountof iron that does dissolve complexes with the N2O4 to form[N2O+

3 ] [Fe(NO3)−4 ]. This adduct can precipitate in valvesand small orifices to cause a phenomenon called “flowdecay.” Users often condition the oxidizer by cooling to40◦F (278 K) and filtering through a 2 µm filter. Subse-quent storage in stainless steel will slow down the for-mation of the “flow decay” adduct, but not eliminate it.Propellant-feed systems of titanium or aluminum are usedin missiles and satellites where mission life approaches20 years.

Pump transfer is recommended for nitrogen tetroxideand the MONs. Storage and transfer areas must be keptfree of oils, fuels, and combustibles.

5. Transportation

N2O4 and the MONs are transported in specially con-structed semitrailers, rail cars, and in high-pressure, seam-less steel cylinders. Carbon steel equipment can be used,but the trend is to use stainless steel. The DOT classifica-tion is Class A Poison Gas and a white label is required.

J. Otto Fuel II

Otto Fuel II is a very stable monopropellant, developedto provide a safe fuel for use in undersea weapons. It isa bright red, oily liquid, insoluble in water but is solublein most common solvents. Otto Fuel is thermally stable at122◦F (323 K). Very slow decomposition may occur above150◦F (339 K), and above 290◦F (416 K) rapid exothermicdecomposition leading to container rupture and fire willoccur.

1. Specifications

Otto Fuel II is composed of a desensitized nitrate ester.The composition of Otto Fuel II is given in Table XV, andis covered by Specification MIL-O-82672(OS).


Most common metals, except those containing copper, aresuitable for Otto Fuel service. The following nonmetalsare satisfactory for contact with Otto Fuel: Butyl rub-ber: ethylenepropylene rubber; glass; polyethylene; poly-tetrafluoroethylene; polychlorotrifluoroethylene; and nat-ural rubber. Buna-N, Neopren

❤™, and Viton❤™ will swell in

contact with Otto Fuel. Polysulfide rubber is incompati-ble. Most fluorocarbon and silicon-based oils and greasesare acceptable. Petroleum-based lubricants are not recom-mended.

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This propellant offers outstanding safety characteristics.Under atmospheric conditions, it is considered to be non-flammable and nonexplosive, making it safe to handle withproper precautions. The propellant has a high flash pointof 265◦F (403 K), permitting it to be classified as a low firehazard material. It is noncorrosive. The vapors are toxic,with a TLV of 0.2 ppm; however, because of the extremelylow vapor pressure it normally does not present a toxicityproblem. Toxic effects may occur from ingestion and skincontact.


Typically 6000-gal (22.7 m3) tanks are used for bulk stor-age of Otto Fuel. The 55-gal (0.2 m3) shipping drums maybe used for storage of small quantities. The containersmust be equipped with pressure relief devices to preventinternal pressure buildup over 60 psi (414 kN/m2).

5. Transportation

Otto Fuel is shipped by 5000 gal (19 m3) tank trucks,preferably made of stainless steel. Polyethylene 55-gal(0.2 m3) drums overpacked with a steel drum are usedfor small shipments. An ullage of 5% is required. Allcontainers must be provided with a means to vent excesspressure.

K. Oxygen

Oxygen is a strong oxidizer. Liquefied, highpurity oxygenis a light-blue transparent liquid that boils vigorously atnormal conditions and has no odor. It is not shock sensitive,does not decompose, and is chemically stable. Liquefiedoxygen is a widely used commercial product.

1. Specifications

Propellant-grade liquefied oxygen contains a minimum of99.5 wt% O2 with the major impurity being argon. Theprime requirement for oxygen purity is based on safety.Hydrocarbons in oxygen produce a very hazardous con-dition. Hydrocarbon impurities that have low solubilitiesare especially undesirable since they are usually frozensolid and can act as ignition sources. Acetylene is lim-ited to 0.5 ppm maximum for this reason. The militaryspecification is MIL-P-25508.


When selecting materials for liquefied oxygen, consider-ation should be given to low temperature properties. The

ability to withstand stress resulting from rapid temperaturechanges is especially important. The following metals aresuitable for liquefied oxygen service: stainless steel—304,304L, 304ELC, 310, 316, 321; aluminum alloys—1000,2014, 2024, 3000, 5050, 5052, 5083, 5154, 6061, 6063,7075; copper and copper alloys—copper, naval and ad-miralty brass, aluminum bronze, cupro-nickel; and nickeland nickel alloys—nickel, 9% nickel steel alloy,InconelX

❤™, Hastelloy❤™ B, Rene 41, K-Monel

❤™. Titanium6A14V can be used in some flight systems, but it is bestavoided in storage or test facilities because of its lowignition temperature in oxygen. In fact, metals will ignitein pure oxygen at high enough pressure or high enoughflow velocity.

Most nonmetal materials lose their elastomeric proper-ties or flexibility at low temperatures, thus limiting thenumber of materials suitable for liquefied oxygen ser-vice. The following have been used successfully: tetraflu-oroethylene polymer (Teflon

❤™); polychlorotrifluoroethy-lene, unplasticized (Kel-F

❤™); asbestos, oil and grease free;and special low-temperature silicon rubber.

Specialized lubricants such as KRYTOX❤™ and thread

sealants are available. The fluorolubes and perfluorocar-bon based products must not contain hydrocarbons, andshould not be used in aluminum equipment.


The principal hazards associated with the use of liquefiedoxygen are

1. Frostbite when the liquid, or uninsulated pipingcontaining it, contacts the skin

2. Fire supported by oxygen3. Formation of shock-sensitive mixtures, especially

with liquid fuels

There are no toxic effects. Oxygen does not burn, butvigorously supports combustion. It is hypergolic with py-rophoric materials, such as triethylaluminum, but not withcommon rocket fuels. Gaseous oxygen resulting fromliquid vaporization adsorbs readily in clothing, and anysource of ignition may cause flare burning.


Liquefied oxygen should be stored in stationary, portable,or mobile tanks designed specifically for oxygen service.Both storage and transportable containers should be vac-uum jacketed; the vacuum may contain reflective insula-tion or noncombustible powders. Either an open vent ora pressure-relieving device to permit the escape of vaporis required. A warm vessel should be filled slowly until

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the system reaches an equilibrium temperature to reducethermal shock. Considerable boil-off of oxygen will occurduring an initial fill and must be vented away from flames,combustibles, and heated surfaces.

5. Transportation

Rail tank cars and tank trucks are used for shipping lique-fied oxygen. Shipments at atmospheric pressure, or under25 psig (172.4 kN/m2) are not regulated; however, pres-sures over 25 psig (172.4 kN/m2) are defined by DOT asa nonflammable compressed gas.

SEE ALSO THE FOLLOWING ARTICLES

COMBUSTION • FUELS • JET AND GAS TURBINE ENGINES

• RAMJETS AND SCRAMJETS • ROCKET MOTORS, HY-BRID • ROCKET MOTORS, LIQUID • ROCKET MOTORS,SOLID • SOLID PROPELLANTS • SPACECRAFT CHEMICAL

PROPULSION

BIBLIOGRAPHY

Clark, J. D. (1972). “Ignition!” Rutgers Univ Press, New Brunswick,New Jersey.

Holzmann, R. T. (1969). “Chemical Rockets & Flame and ExplosivesTechnology,” Dekker, New York.

Glassman, I., and Sawyer, R. F. (1970). “The Performance of Chemi-cal Propellants,” AGARDograph 129, Technivision Services, Slough,England.

Palaszewski, B. (1997). “Propellant Technologies: A Persuasive Waveof Future Propulsion Benefits,” February.

Sarner, S. F. (1966). “Propellant Chemistry,” Reinhold, New York.Schmidt, E. W. (1984). “Hydrazine and Its Derivatives,” Wiley, New

York.Stull, D. R., Prohpet, H., et al. (1971). JANAF Thermochemical Tables,

Second Edition. NSRDS-NBS-37, SD Catalog No. C13.48.37. Officeof Standard Reference Data, National Bureau of Standards, Washing-ton, DC.

Sutton, George P. (1986). “Rocket Propulsion Elements,” 5th edition.Wiley, New York.

Hazards of Chemical Rockets and Propellants (1984). Vol III, “Liq-uid Propellants,” CPIA Publ. 394, Chemical Propulsion InformationAgency, Laurel, Maryland.

United States Patent 5,616,882. High Energy Rocket Propellant, April 1,1977.

Weber, J. Q. (1984). Advanced Liquid Propellant Systems for ChemicalPropulsion, Progr. Astronaut. Aeronaut. 89, 559–568.

Web Siteshttp://members.aol.com/ensave/astro.html X-L Space Systemshttp://www.lerc.nasa.gov NASA Lewis Research Centerhttp://www1.msfc.nasa.gov/NEWSROOM NASA Marshall Space

Flight Centerhttp://www.grc.nasa.gov NASA Langley Research Centerhttp://www.afrl.af.mil/dualuse/affas.htm Air Force Research Lab, Pro-

pulsion Directoratehttp://www.chem.utah.edu/chemistryhttp://www.personal.psu.eduhttp://www.sigma-aldrich.com

%E2%80%A2%20%E2%80%A2%20PROPULSION%0D%0Ahttp://members.aol.com/ensave/astro.html

http://www.lerc.nasa.gov

http://www1.msfc.nasa.gov/NEWSROOM

http://www.grc.nasa.gov

http://www.afrl.af.mil/dualuse/affas.htm

http://www.chem.utah.edu/chemistry

http://www.personal.psu.edu

http://www.sigma-aldrich.com

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